(iv) Bearing Accuracy. Using a properly operating directional gyro, and a ground checkpoint whose location is known with respect to the ground-based station to which the ADF is tuned, fly the airplane over the ground checkpoint at a minimum of six airplane headings (relative to the station to which the ADF is tuned), including headings of 0 degrees, 180 degrees, and headings 15 degrees each side thereof. Determine that the compensated or otherwise corrected ADF bearing indication is not in error by more than ±5 degrees at any of the six headings. Repeat the foregoing procedure but fly at a minimum of six headings relative to the ground based station, including 45 degrees, 90 degrees, 135 degrees, 225 degrees, 270 degrees, and 315 degrees. In this case, the compensated or otherwise corrected bearing indication should not be in error by more than ±10 degrees at any of these six headings.
NOTE: The distance between the ground checkpoint and the ground-based station used should be at least one-half of the service range of that station.
(v) Indicator Response. With the ADF indicating station dead ahead, switch to a station having a relative bearing of 175 degrees. The indicator should indicate within ±3 degrees of the bearing in not more than 10 seconds.
(vi) Antenna Mutual Interaction. If the ADF installation being tested is dual, check for coupling between the antennae by using the following procedures.
(A) With #1 ADF receiver tuned to a station near the low end of the ADF band, tune the #2 receiver slowly throughout the frequency range of all bands and determine whether the #1 ADF indicator is adversely affected.
(B) Repeat (A) with the #1 ADF receiver tuned to a station near the high end of the ADF band.
(vii) Electro-Magnetic Compatibility (EMC). With all systems operating in flight, verify by observation that no adverse effects are present in the required flight systems.
(viii) Precipitation Static. The effect of precipitation static should be considered. This type of static is normally found in areas of high cirrus clouds, dry snow, dust storms, etc.
(6) Distance Measuring Equipment (DME) .
(i) Airplanes to be operated above 18,000 ft. The DME system should continue to track without dropouts when the airplane is maneuvered throughout the airspace within 160 NM of the VORTAC station and at altitudes above radio line-of-sight up to the maximum altitude for which the airplane is certificated. This tracking standard should be met with the airplane:
(A) In cruise configuration.
(B) At bank angles up to 10 degrees (skidding turns may be used to minimize turn radius).
(C) Climbing and descending at n ormal maximum climb and descent attitudes.
(D) Orbiting a DME facility.
(E) Providing clearly intelligible (visual/aural) identification of the DME facility.
(ii) Airplanes to be operated at altitudes below 18,000 ft. The DME system should perform as specified in paragraph ( i) above, except that the maximum distance from the DME facility need not exceed 80 NM.
(iii) Climb and Maximum Distance. Determine that there is no mutual interference between the DME system and other equipment aboard the airplane. Beginning at a distance of at least 10 NM from a DME facility and at an altitude of 2,000 ft. above the DME facility, fly the airplane on a heading so it will pass over the facility. At a distance of 5 to 10 NM beyond the DME facility, operate the airplane at its normal maximum climb attitude, up to an altitude of 35,000 ft. or to within 90 percent of the maximum certificated altitude, maintaining the
airplane on a station radial (within 5 degrees). The DME should continue to track with no unlocks to a range of 160 NM (80 NM for airplanes not to be operated at altitudes above 18,000 ft.).
(iv) Long Range Reception. Perform two 360 degree turns, one to the right and one to the left, at a bank angle of 8 to 10 degrees at not less than 160 NM (80 NM for airplanes not to be operated above 18,000 ft.)
from the DME facility. A single turn will be sufficient if the antenna installation is symmetrical. There should be no more than one unlock, not to exceed one search cycle (maximum 35 seconds) in any five miles of radial flight.
Tests may be conducted up to a maximum certificated altitude.
(v) High Angle Reception. Repeat the flight pattern and observations of (iv) above at a distance of 50-70 NM (20 to 30 NM for airplanes not to be operated above 18,000 ft.) from the DME facility and at an altitude of at least 35,000 ft. (or to within 90 percent of the maximum altitude for which the airplane is certificated or the maximum operating altitude).
(vi) Penetration. From an altitude of at least 35,000 ft. (or the airplane maximum altitude, if lower), perform a letdown directly toward the ground station using normal maximum rate of descent procedures to a DME facility, so as to reach an altitude of 5,000 ft. above the DME facility 5 to 10 NM before reaching the DME facility. The DME should continue to track during the maneuver with no unlocks.
(vii) Orbiting. At an altitude of 2,000 ft. above the terrain, at holding pattern speeds appropriate for the type of airplane and with the landing gear extended, fly at least 15 degree sectors of left and right 35 NM orbital patterns around the DME facility. The DME should continue to track with no more than one unlock, not to exceed one search cycle, in any 5 miles of orbited flight.
(viii) Approach. Make a normal approach to a field with a DME. The DME should track without an unlock (station passage excepted).
(ix) Electro-Magnetic Compatibility (EMC). With all systems operating in flight, verify by observation that no adverse effects are present in the required flight systems.
(7) Transponder Equipment .
(i) The ATC transponder system should furnish a strong and stable return signal to the interrogating radar facility when the airplane is flown in straight and level flight throughout the airspace within 160 NM of the radar station, from radio line-of-sight to within 90 percent of the maximum altitude for which the airplane is certificated or to the maximum operating altitude. Airplanes to be operated at altitudes not exceeding 18,000 ft.
should meet the above requirements to only 80 NM.
(ii) When the airplane is flown in the following maneuvers within the airspace described above, the dropout time should not exceed 36 seconds:
(A) In turns at bank angles up to 10 degrees (skidding turns may be used to minimize turn radius).
(B) Climbing and descending at typical climb and descent attitudes.
(iii) Climb and Distance Coverage.
(A) Beginning at a distance of at least 10 NM from, and at an altitude of 2, 000 ft. above that of the radar facility, and using a transponder code assigned by the Air Route Traffic Control Center (ARTCC), fly on a heading that will pass the airplane over the facility. Operate the airplane at its normal maximum climb attitude up to within 90 percent of the maximum altitude for which the airplane is certificated, or to a maximum operating altitude, maintaining the airplane at a heading within 5 degrees from the radar facility. After reaching the maximum altitude for which the airplane is certificated, fly level at the maximum altitude to 160 NM (or 80 NM for airplanes to be operated below 18,000 ft.) from the radar facility.
(B) Communicate with the ground radar personnel for evidence of transponder dropout.
During the flight, check the " ident" mode of the ATC transponder to assure that it is performing its intended
function. Determine that the transponder system does not interfere with other systems aboard the airplane and that other equipment does not interfere with the operation of the transponder system. There should be no dropouts, that is, when there is no return for two or more sweeps. If ring-around, spoking, or clutter appears on the ground radar scope, the airplane should switch to "low" sensitivity to reduce the interference. Uncontrollable ringing that hinders use of the ground radar should be considered unsatisfactory.
(iv) Long Range Reception. At 90 percent of maximum certificated altitude, perform two 360 degree turns, one to the right and one to the left, at bank angles of 8 degrees and 10 degrees with the flight pattern at least 160 NM (or 80 NM for airplanes to be operated below 18,000 ft.) from the radar facility. During these turns, the radar display should be monitored and there should be no signal dropouts (two or more sweeps). (Airspeed adjustments may be made to reduce turn radius.)
(v) High Angle Reception. Repeat the flight pattern and observations of (iv), above, at a distance of 50 to 70 NM from the radar facility and at an altitude of at least 35,000 ft. or within 90 percent of the maximum altitude for which the airplane is certificated or the maximum operating altitude. There should be no dropout (two or more sweeps). Switch the transponder to a code not selected by the ground controller. The airplane secondary return should disappear from the scope. The controller should then change his control box to a common system and a single slash should appear on the scope at the airplane's position. If a problem surfaces, a placard so noting shall be required.
(vi) High Altitude Cruise. Fly the airplane within 90 percent of its maximum certificated altitude or its maximum operating altitude beginning at a point 160 NM (or 80 NM for airplanes to be operated below 18,000 ft.) from the radar facility on a course that will pass over the radar facility. There should be no transponder dropouts (two or more sweeps) or "ring-around," except when station passage occurs or when known site features produce anomalies.
(vii) Holding and Orbitin g Patterns.
(A) At an altitude of 2,000 feet or minimum obstruction clearance altitude (whichever is greater) above the radar antenna, and at holding pattern speeds, flaps and gear extended, fly one each standard rate 360 degree turn, right and left, at a distance of approximately 10 NM from the Air Route Surveillance Radar (ARSR) or Airport Surveillance Radar (ASR) facility. There should be no signal dropout (two or more sweeps).
(B) At an altitude of 2,000 feet or minimum obstruction clearance altitude (whichever is greater) above the radar antenna and at holding pattern speeds appropriate for the type of airplane, fly a 45 degree segment of a 10 NM orbit centered on the radar facility with gear and flaps extended. There should be no signal dropout (two or more sweeps).
(viii) Surveillance Approach. From an altitude of 35,000 feet or within 90 percent of the maximum certificated altitude of the airplane or the maximum operating altitude, whichever is less, perform a letdown and approach to a runway of an airport served by ASR having an Air Traffic Control Radar Beacon System (ATCRBS) facility. Alternately, a simulated approach and letdown may be made along a path parallel to, but separated three to four miles from, a vertical plane through the location of an ARSR facility. The approach should be made at the maximum normal rate of descent and in the normal approach and landing configuration for the airplane, and should continue down to an altitude of 200 feet or less above the ground radar antenna elevation. Not more than one dropout should occur for any 10 sweeps during final approach; if ring-around occurs, the flightcrew should reduce the sensitivity and the ring-around should cease.
(ix) Altitude Reporting. Conduct a functional tes t of the altitude encoder (transponder Mode "C") by comparison with ATC-displayed altitudes. Verify correspondence at several altitudes between ATC readings and the captain's altimeter, when set at or corrected to 29.92 inches of mercury.
NOTE: Throughout all tests, verify self-test function and radar interrogation reply light.
(x) Electro-Magnetic Compatibility (EMC). With all systems operating in flight, verify by observation that no adverse effects are present in the required flight systems.
(8) Weather Radar System .
(i) Warm-up Period (if applicable.) All tests shall be conducted after the manufacturer's specified warm-up period.
(ii) Display. Check that the scope trace and sweep display move smoothly and are without gaps or objectionable variations in intensity. For color displays, verify appropriate colors and color contrast.
(iii) Range Capabilities. While maintaining level flight at 90 percent of maximum approved altitude of the airplane, set the radar controls so that large radar-identifiable objects such as mountains, lakes, rivers, coastlines, storm fronts, turbulence, etc., are displayed on the radar scope. Objects should be displayed at the maximum range specified by the manufacturer. Maneuver the airplane and adjust the radar controls so that tests may be conducted for the range requirements. The radar should be capable of displaying line-of-sight known prominent targets.
(iv) Beam Tilting and Structural Clearance.
(A) With the airplane maintained in level fl ight, adjust the radar controls so that large radaridentifiable targets appear on the radar scope when the antenna tilt control is adjusted for zero degrees. Maneuver the airplane so that an appropriate target appears at the dead ahead (0 degrees) bearing position. Slowly change the tilt control through an appropriate range and observe that the radar scope presentation does not change erratically, which might indicate structural interference between the radome and antenna. This interference should not occur throughout the operational speed envelope of the airplane.
(B) Bearing Accuracy. Fly under conditions that allow visual identification of a target, such as an island, a river, or a lake, at a range within the range of the radar. When flying toward the target, select a course from the reference point to the target and determine the error in the displayed bearing to the target on all range settings. Change the heading of the airplane in increments of 10 degrees to ±30 degrees maximum and verify the error does not exceed ±5 degrees maximum in the displayed bearing to the target.
(v) Stability. While observing a target return on the radar indicator, turn off the stabilizing function and put the airplane through pitch and roll movements. Observe the blurring of the display. Turn the stabilizing mechanism on and repeat the roll and pitch movements. Evaluate the effectiveness of the stabilizing function in maintaining a sharp display.
(vi) Contour Display ( Iso Echo).
(A) If heavy cloud formations or rainstorms are reported within a reasonable distance from the test base, select the contour display mode. The radar should differentiate between heavy and light precipitation.
(B) In the absence of the above weather conditions, determi ne the effectiveness of the contour display function by switching from normal to contour display while observing large objects of varying brightness on the indicator. The brightest object should become the darkest when switching from normal to contour mode.
(vii) Antenna Stability When Installed. While in level flight at 10,000 feet or higher, adjust the tilt approximately 2-3 degrees above the point where ground return was eliminated. Put the airplane through down pitch, then roll movements. No ground return should be present.
NOTE: Roll right and left approximately 15 degrees; pitch nose down approximately 10 degrees.
(viii) Ground Mapping. Fly over areas containing large, easily identifiable landmarks such as rivers, towns, islands, coastlines, etc. Compare the form of these objects on the indicator with their actual shape as visually observed from the cockpit.
(ix) Mutual Interference. Determine that no objectionable interference is present on the radar indicator from any electrical or radio/navigational equipment, when operating, and that the radar installation does not interfere with the operation of any other equipment. Particular attention should be devoted to radar effect on electronic display systems and on localizer, glideslope, and ADF signals during approach to landing.
(x) Electro-Magnetic Compatibility (EMC). With all systems operating in flight, verify by observation that no adverse effects are present in the required flight systems.
(9) Inertial Navigation System (INS) .
(i) Advisory Circular 25-4, "Inertial Navigation System," dated February 16, 1966, contains the basic criteria for the engineering evaluation of an inertial navigation system and offers acceptable means of compliance with the applicable regulations that contain mandatory requirements in an objective form. The engineering evaluation of an INS should also include awareness of Advisory Circular 121-13, " Self-Contained Navigation Systems (Long Range)," dated October 14, 1969, which presents criteria to be met before an operator can obtain operational approval. For flights up to 10 hours, the radial error should not exceed 2 nautical miles per hour of operation on a 95 percent statistical basis. For flights longer than 10 hours, the error should not exceed ±20 NM cross-track or ±25 NM along-track error. A 2 nautical mile radial error is represented by a circle, having a radius of 2 nautical miles, centered on the selected destination point. For airplanes intended for air carrier operation, reference should be made to Appendix G to Part 121 of the FAR.
(ii) The INS equipment supplier, through the applicant, will generally provide system reliability data sufficient so that additional testing for equipment reliability demonstration will not be required. Assuming adequate cooling is available for the installation, the data can be considered independent of which type of airplane was used for derivation.
(iii) Temperature. Some systems have temperature monitors built into the sensor block. Whe n the device temperature reaches a given level, the system automatically "shuts-down." This condition could represent a common mode failure wherein insufficient cooling is provided to the multiple sensors with the consequent result that they all "trip" the temperature monitor simultaneously. Some equipment may be constructed with the cooling mechanization integral to the individual unit. Regardless of how the equipment cooling is accomplished, if the proper operation of the unit is below acceptable levels due to failures of the cooling function, then the cooling function must be addressed by analysis and demonstration, where applicable.
(iv) INS On-Airplane Test Procedure.
(A) With INS equipment off, verify all avionics equipment are operating n ormally.
(B) Close INS primary, heater, and excitation circuit breakers and set the mode selector to start warm-up, as required by the manufacturer's normal procedures.
(C) Alignment.
(1) Initiate INS alignment according to the manufacturer 's procedures and input flight plan waypoints and record. (The airplane's present position must be loaded before the system can complete alignment. The airplane must not be moved until the system is aligned and in the "Navigate" mode, if so equipped.)
Present position accuracy should be known to within 0.1 NM.
(2) Normal wind buffeting and movement of cargo and personnel will not affect alignment quality, but severe buffeting may lengthen the time required for alignment.
(3) Continue operatio n in the “NAV” mode until a minimum of one hour has passed.
(D) Pre-departure.
(1) Recheck that other avionics operate normally and that all systems and instruments using INS information operate correctly.
(2) Set up INS for a track from p resent position to any stored waypoint and check proper cockpit display unit (CDU) and instrument displays.
(3) When the conditions as described by the manufacturer are satisfied, set the mode selector to "Navigate". Verify that the NAV mode is achieved and no "warn" condition is indicated. Record the time at which NAV mode is verified.
(E) Departure Procedure.
(1) Taxi. While taxiing the airplane, monitor INS performance with the CDU data selector set to display track angle and ground speed.
(2) Takeoff. With the CDU set as required to observe ground speed, track angle, HDG, and drift angle during takeoff and initial airborne conditions, verify normal and proper performance. Reasonable assessment may be accomplished using other instruments and avionics..
(3) Initial track. Define a track from present position to first available waypoint and fly to intercept using instrument display or autopilot.
(F) In Flight. During flight conditions, perform and/or check the followi ng: (NOTE: Do not change system mode selector from "Navigate" during navigational tests. If system is installed to provide attitude for the airplane system, the "ATT REF" mode must be checked after navigation mode tests are completed). When selector is changed to "ATT REF," ascertain that all appropriate equipment annunciators indicate the disconnect.)
(1) Exercise all procedures, such as waypoint navigation, offset navigation, selected track navigation, etc., that are listed as system capabilities by the manufacturer and in the Airplane Flight Manual (AFM).
(2) If updating capabilities are provided, either as manual or automatic, use defined procedures to verify operation.
(3) Operate the INS in both the manual and automatic modes for waypoint sequencing.
(4) Verify satisfactory control of the airplane, proper flight director indication, and flight instrument presentation, as appropriate to the installation. When the INS is operating, all airplane displays that would be affected or changed in meaning should be clearly annunciated to the pilot (e.g., compass heading (true versus magnetic) autopilot coupled modes, etc.).
(5) Operate the CDU in all display settings and verify functions and operation.
(6) Operate other avi onics and airplane electrical systems to determine if any mutual interference results that would cause either the INS or other systems' performance to be less than required.
(7) If basic accuracy has been previously established, a triangular course s hould be flown with a base leg length of at least one hour. The required accuracy is defined in paragraph 170(9)( i), above.
(G) Post Flight.
(1) Park the airplane at a position known to within 0.1 NM and determine the radial error rate using the total time in NAV mode and the measured radial error. Record end point position, time, and error rate calculations.
NOTE 1: On some models, radial error may be obtained as follows: Load the park position as waypoint N; then define track (LEG) O to N. The value of distance read on the CDU is the radial error.
NOTE 2: If the system has been updated, remove the update prior to error calculations.
(2) Pull INS primary power circuit breaker and verify operation on battery for 5 minutes. Reset breaker and verify that the system functions normally and the waypoint data is correct.
(3) Turn system to off.
(v) Attitude (if provided to the basic instruments).
(A) The system must be installed to provide pitch and roll data to ±1 degree re lative to the airplane's reference level.
(B) During navigation tests, the vertical gyro data should be checked. Additionally after NAV tests are completed, the airplane should be flown and placed in a 25 degree bank with the INS in NAV mode.
While in the turn, switch to ATT REF mode, and level the airplane. The attitude data should indicate properly. If the system requires level flight at the time ATT REF is selected, the AFM should contain the limitation.
(C) Power Bus Transients. After the normally expected electrical bus transients due to engine failure, attitude should not be off or unstable for more than one second and should affect only displays on one side of the airplane. If power-up initialization or self-tests are started by the transient, any change in attitude should not be a distraction; recognizably valid pitch and roll data should be available within one second. For most airplanes, an engine failure after takeoff will simultaneously create a roll rate acceleration, new pitch attitude requirements, and an electrical transient. Attitude information is paramount; transfer to standby attitude or transfer of control of the airplane to the opposite pilot cannot be reliably accomplished under these conditions in a timely enough fashion to prevent an unsafe condition. In testing this failure mode, switching the generator off at the control panel will usually result in the quickest switching time. Conversely, during an engine failure, as the engine speed decays, the generator output voltage and frequency each decay to a point where the bus control relays finally recognize the failure. This can be a significantly larger disturbance resulting in a different effect on the using equipment. The only known way to simulate this failure is with a fuel cut. Both means should be tested.
(vi) Electromagnetic Compatibility. The INS should not cause the performance of other systems aboard the airplane to be degraded below their normal function, and INS operation should not be adversely affected by other equipment.
(10) Omega/VLF Navigation System . A complete and comprehensive flight test evaluation is contained in Advisory Circular 20-101C, "Airworthiness Approval of Omega/VLF Navigation Systems for use in the U.S.
National Airspace System (NAS) and Alaska,” dated September 12, 1988.
(11) Loran-C Navigation System . A complete and comprehensive flight test evaluation is contained in Advisory Circular 20-121A, "Airworthiness Approval of Loran-C Navigation Systems for use in the U.S. National Airspace System (NAS) And Alaska," dated August 24, 1988.
(12) Doppler Navigation System . Doppler Navigation System installed performance should be evaluated in accordance with Advisory Circular 121-13, "Self Contained Navigation Systems (Long Range)," dated October 14, 1969. (See also 14 CFR part 121, Appendix G, and paragraph 170c(9) of this AC, which addresses Inertial Navigation Systems.)
(13) Area Navigation (RNAV) . Information pertinent to flight evaluation of area navigation is contained in Advisory Circular 90-45A, "Approval of Area Navigation Systems for Use in the U.S. National Airspace System," dated February 21, 1975. Note that the area navigation function is most often performed by the systems identified in paragraphs 170c(9), (10), (11), (12), and (14) of this AC. Flight test requirements should be reviewed accordingly.
(14) Multi-Sensor Navigation System . A complete and comprehensive flight test evaluation is contained in Advisory Circular 20-130A, "Airworthiness Approval of Navigation or Flight Management Systems Integrating Multiple Navigation Sensors,” dated June 14, 1995.
(15) Vertical Navigation System (VNAV) . A complete and comprehensive flight test evaluation is contained in Advisory Circular 20-129 "Airworthiness Approval Of Vertical Navigation (VNAV) Systems for use in the U.S. National Airspace (NAS) and Alaska," dated September 12, 1988.
(16) Performance Management System (PMS) . A complete and comprehensive flight test evaluation is contained in Advisory Circular 25-15 "Approval of Flight Management Systems In Transport Category Airplanes," dated November 20, 1989.
(17) Flight Management System (FMS) . A complete and comprehensive flight test evaluation is contained in Advisory Circular 25-15, "Approval of Flight Management Systems In Transport Category Airplanes," dated November 20, 1989.
(18) Global Positioning System (GPS) . A complete and comprehensive flight test evaluation is contained in Advisory Circular 20-138, “Airworthiness Approval Of Global Positioning Systems (GPS) Navigation Equipment For Use As A VFR And IFR Supplemental Navigation System,” dated May 25, 1994.
(19) Microwave Landing System (MLS) . [RESERVED]
d. Procedures-Instruments and Displays .
(1) Flight and Navigation Instruments .
(i) Flight demonstration of the following pilots' instruments should address evaluation with regard to display, function, and lighting:
(A) Attitude Director Indicator (ADI)-primary and standby (B) Horizontal Situation Indi cator (HSI) (C) Radio Magnetic Indicator (RMI) (D) Airspeed and Mach No. Indicator-primary and standby (if installed) (E) Altimeter-primary and standby (F) Instantaneous Vertical Speed Indicator (IVSI) (G) Standby Compass (H) Clock (I) Total Air Temperature Indicator (TAT) (J) Radio Altimeter Indicator (R/A)
(ii) The instruments should be evaluated for suitability of location and performance during flight operation. Agreement between displayed functions, primary to standby, and captain's to first officer's display (where applicable) should be examined during takeoff, cruise, and landing, and during flight maneuvers of bank angles of ±60 degrees and pitch angles from + 25 to -10 degrees.
(iii) In addition to the above instruments, the controls on the instrument, and those associated with changing the function of that instrument, should be evaluated.
(iv) Day and night lighting of instruments must be demonstrated. Evaluate for adequate illumination of the instruments and associated control panels and placards.
(2) Electronic Display Systems . A complete and comprehensive description and flight evaluation is contained in Advisory Circular 25-11 "Transport Category Airplane Electronic Display Systems," dated July 16, 1987.
e. Procedures-Sensors and Warning Systems .
(1) Compass System .
(i) Conduct a ground compass swing and record deviations from correct magnetic headings at 15 degree intervals. Be particularly attentive for errors induced by intermittently operated electrical equipment that is not removable by compass compensation plates. (See chapter 16, section 5, of AC 43.13-1A, “Acceptable Methods, Techniques and Practices,” dated 1988, or latest revision.)
(ii) Compare indicated heading information with known runway directions during takeoffs and landings.
(iii) Conduct an inflight compass check by stabilizing the airplane on the four cardinal headings and recording the respective compass headings of both primary compass systems (in the slaved mode) and the
standby compass system. These readings should be compared with the flight recorder data as a check on compass repeater capability.
(iv) Observe indicators for evidence of electronic or magnetic interference. Specifically, observe the standby compass operation as a function of windshield heat cycling and while keying radio transmissions.
(2) Attitude System-"Strapdown" (AHRS & IRS) .
(i) In addition to flight evaluation of the pitch and roll attitude characteristics of conventional vertical gyros, which is rather straight-forward, the microprocessor-based strapdown attitude (and heading) systems present some additional considerations for certification and flight testing, which are discussed herein. Attitude and Heading Reference Systems (AHRS) and Inertial Reference Systems (IRS) are the types of systems envisioned.
(ii) Validation of acceptable equipment installations includes, but is not limited to, the validation of proper installation considering the combined effects of temperature, altitude, EMI, vibration, and other various environmental influences. These installation requirements are applicable to critical, essential, and non-essential systems. However, there may be cases where non-essential installations do not warrant the expense of having all, or even part, of these tests and analyses conducted. The necessity for conducting these on non-essential installations should, therefore, be determined on a case-by-case basis by the FAA project engineer, based on the specific and individual circumstances involved. Particular attention should be paid to the following environmental considerations:
(A) Vibration. Testing is generally accomplished on the sensor level to DO-160C requirements. The structural mounting provisions on the airplane should provide assurance that the DO-160C levels are not exceeded when the airplane encounters gusts, as defined in Appendix G to Part 25. The concern is that the dynamic range of the sensor may be exceeded during this vibration encounter.
(B) Temperature. Some systems have temperature monitors built into the sensor block.
When the device temperature reaches a given level, the system automatically "shuts-down." This condition could represent a common mode failure wherein insufficient cooling is provided to the multiple sensors, with the consequent result that they all "trip" the temperature monitor simultaneously. Some equipment may be constructed with the cooling mechanization integral to the individual unit. Regardless of how the equipment cooling is accomplished, if the proper operation of the unit is below acceptable levels due to failures of the cooling function, then the cooling function must be addressed by analysis and demonstration, where applicable.
(C) Power Bus Transients. After the normally expected electrical bus transients due to engine failure, attitude should not be off or unstable for more than one second and should affect only displays on one side of the airplane. If power-up initialization or self-tests are started by the transient, any change in attitude should not be distracting; recognizably valid pitch and roll data should be available within one second. For most airplanes, an engine failure after takeoff will simultaneously create a roll rate acceleration, new pitch attitude requirements, and an electrical transient. Attitude information is paramount; transfer to standby attitude or transfer of control of the airplane to the opposite pilot cannot be reliably accomplished under these conditions in a timely enough fashion to prevent an unsafe condition. In testing this failure mode, switching the generator off at the control panel will usually result in the quickest switching time. Conversely, during an engine failure, as the engine speed decays, the generator output voltage and frequency each decay to a point where the bus control relays finally recognize the failure. This can be a significantly larger disturbance resulting in a different effect on the using equipment. One way to simulate this failure is with a fuel cut, at a safe altitude and airspeed. Both means should be tested.
(iii) Multi-axis Failures - Failure Modes and Effects Analyses ( FMEAs) generally do not rule out the combined pitch/roll attitude failures. If this failure condition can propagate through the autopilot, a need may exist for demonstrating multi-axis autopilot hardovers, as described in paragraph 181 of this AC. If the autopilot system architecture is such that the failure cannot get through to the control surface (i.e., input sensor screening, failpassive autopilot, reasonableness tests, etc.), the need for hardover demonstration may be eliminated. Be aware of the fact that most fail-passive cruise mode autopilots may not be truly fail-passive when reconfigured to reversionary modes.
(iv) Attitude Sensor Intermix - If the applicant is seeking certification credit for sensor intermix
(e.g., vertical and directional gyros (VG/DG) on the captain's side, and attitude/heading reference system (AHRS)
on the first officer's side), compatibility of the two systems should be demonstrated (i.e., no excessive attitude monitor trips, no adverse influence on system availability (such as autoland) due to monitor trips, etc.). This also holds true for either mode of AHRS operation (NORMAL/BASIC or STANDBY), if this feature is available in the system.
(v) For systems that have two operational modes (NORMAL and BASIC), annunciation of the fact that the system is operating in the reversionary (BASIC) mode is largely dependent on the "user" system's architecture. For example, if the airplane is equipped with a CAT III autoland system or CAT III Head-up display (HUD) system, the performance in these modes with Basic Attitude may be degraded to the point where airplane safety is compromised in CAT III weather conditions. There must then be clear and unmistakable annunciation of the fact that this attitude condition exists.
(vi) For triple AHRS installations, switching is provided to substitute the third AHRS for either the captain's or first officer's system. Annunciation of the fact that the switched condition is in existence may or may not be required, depending on cockpit layout, pilot workload considerations, etc. This is an assessment generally made by the FAA project pilot. The same holds true for annunciation of whether or not the third system is inoperative.
(vii) Some additional flight evaluation suggestions follow:
(A) While on the ground, determine that the attitude system provides usable attitude information through 360 degrees of roll and pitch by rotating the platform through 360 degrees of roll and pitch while observing the appropriate attitude indicator.
(B) Verify the system performs as intended, providing satisfactory attitude and heading information to the pilot's and copilot's attitude and heading instruments throughout the normal airplane flight envelope, including unusual attitudes that may be expected in service.
(C) Verify that the flight control system functions properly when interfaced with the attitude system. This functional evaluation may be accomplished during a typical flight profile encompassing enroute, maneuvering, and coupled approach operations.
(D) Determine that loss of the air data system or airspeed input is properly annunciated and that the systems continue to provide satisfactory attitude/heading information.
(E) Verify that the comparator monitor provides proper annunciation of attitude/heading disparity.
(F) Verify that loss of a single power source does not cause loss of both attitude systems simultaneously.
(G) Verify that attitude/heading data from each is not lost for more than 1 second following loss of a primary power source (generator, alternator, inverter, etc.) simultaneous with loss of the associated powerplant.
(H) Verify that the system can be realigned in flight after being shutdown for more than 3 minutes (if applicable).
(I) Verify that each attitude system continues to operate correctly following a simulated airplane electrical power loss of "X" seconds (equipment specific), by removing power from each respective electrical bus.
(J) Verify proper operation of the back-up battery (if installed) and display of appropriate flags by pulling the attitude circuit breaker. Note that the system operates correctly on back-up battery power for approximately "X" minutes (equipment specific) and then shuts down. Determine that the appropriate flags appear on the effected displays.
(K) Verify that all controls, displays, and annunciations are satisfactorily identified, accessible, operable, and adequate for direct sunlight and night conditions.
NOTE: It is possible that some of the above items can be accomplished in a laboratory or simulator environment.
(viii) Verify that there is no unacceptable mutual interference between the attitude system and other systems and equipment.
(3) Angle of Attack (AOA) System .
(i) Pilot interface with this system will generally be a consequence of presenting a raw data readout in the cockpit and/or driving the Slow/Fast (S/F) display on the ADI, (if so equipped/)
(ii) Evaluate the displays during normal approaches and landings to assure reasonable/proper information is being presented for the particular flap setting and V REF speed over the gross weight and center of gravity range of the airplane (S/F should be centered).
(iii) Conduct left and right sideslips that would be representative of normal operation on approach during crosswind landings or engine-out conditions.
(iv) Determine that the malfunction indicators (flags) are appropriately displayed and are satisfactory annunciators.
(v) Qualitatively assess that the thresholds of displays are wide enough to permit the pilot to follow the indicators, as appropriate to the operation of the airplane.
(vi) Determine that instrument presentation is sufficiently damped to permit use in turbulent air and that hysteresis, if present, is acceptable.
(vii) Verify that the AOA system neither contributes to, nor is affecte d by, Radio Frequency/Electromagnetic Interference (RFI/EMI).
(4) Air Data System.
(i) The air data system performance, with the exception of airspeed calibration, will generally be demonstrated qualitatively by observation of the output displays.
(ii) Be particularly observant of barometric altimeter and vertical speed reversals as a function of rapid pitch attitude changes (e.g., takeoff rotation).
(iii) Verify that the Mach/Airspeed indicators operate smoothly throughout the complete speed envelope of the airplane.
(iv) Observe correct operation of the overspeed warning indicator (barber pole) while approaching VMO / MMO speeds.
(v) Observe the static and total air temperature (SAT/TAT) indicators for reasonableness of data presentation.
(vi) Verify that the air data system neither contributes to, nor is affected by, Radio Frequency/Electromagnetic Interference.
(5) Radio Altimeter System .
(i) The radio altimeter system should display to the flightcrew, clearly and positively, the altitude information that indicates the airplane main landing gear wheel height above terrain.
(ii) Under the measurement conditions described, verify that the altimeters:
(A) Display altitude without loss of signal indications or excessi ve fluctuations, under the following conditions:
(1) Pitch angle ±5 degrees about the mean approach attitude.
(2) Roll angle zero to ±20 degrees.
(3) Forward velocity from minimum approach speed up to 200 knots, in appropriate configurations.
(4) At altitudes from 0 to 200 feet with sink rates of 0 to 15 feet/second, in landing, approach, and go-around configurations.
(B) Track the actual altitude of the airplane over level ground without significant lag or oscillation.
(iii) With the airplane at an altitude of 200 feet or less, verify that any abrupt change in terrain, representing no more than 10 percent of the airplane's altitude, does not cause the altimeter to unlock. The indicator response to such changes should be appropriate. If the system unlocks, it should re-acquire the signal promptly without pilot intervention.
(iv) Caution should be exercised on airplanes equipped with Automatic Landing Systems if the radio altimeter system is prone to unlock while flying final approach over irregular terrain, since in all probability, the unlock will adversely affect the Autoland Availability requirement.
(v) If a Decision Height (DH) function is provided, verify proper operation at minimum altitudes of 200 feet, 100 feet, and 50 feet.
(vi) Verify that the push-to-test self-test feature generates a simulated radio altitude of less than 500 feet, and that no other systems experience untended effects or interference. For aircraft with automatic landing systems, the Self-Test feature should be shown to be inhibited during the low altitude phases of the autoland.
(vii) Verify that the system provides a positive failure warning display any time there is a loss of power or a failure of the altimeter to function properly.
(viii) Verify that the radio altimeter system neither contributes to, nor is affected by, Radio Frequency or Electromagnetic Interference.
(6) On-Board Weight and Balance System .
(i) In conjunction with conducting the airplane flight test program, on a pre-flight and post-flight basis, continually check and record the gross weight and center of gravity indications of the system with independently and accurately derived values of gross weight and c.g.
(ii) Electromagnetic Compatibility ( EMC). Verify that this system does not create adverse effects on any other required flight systems.
(7) Emergency Locator Transmitters (ELT) . Flight testing of the ELT function is not allowed; however, evaluation of the effects of inadvertent actuation of ELT on other flight systems should be made. One method is to transmit on the VHF Communication System, whose antenna is closest to the ELT antenna, for a period of 3 seconds on frequencies spaced 0.5 Mhz apart throughout the VHF communication band. Verify that the ELT is not activated by any transmission, no adverse interference is generated on VOR systems due to ELT/COM/VOR interactions, and that the desired VHF communication is established.