222 2 In the absence of the above weather conditions, determine 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.
(g) 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.
(h) 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.
(i) 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.
(j) EMC should be evaluated with all systems operating in flight to verify by observation that no adverse effects are present in the required flight systems.
(9) Inertial Navigation System (INS) .
(a) AC 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 AC 121-13, Change 2 “Self-Contained Navigation Systems (Long Range),” dated December 21, 1970, 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 two nautical mile radial error is represented by a circle, having a radius of two 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 CFR.
(b) 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.
(c) 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 should be addressed by analysis and demonstration, where applicable.
(d) INS On-Airplane Test Procedure.
1 With INS equipment off, verify all avionics equipment are operating normally.
2 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.
3 Alignment.
(aa) Initiate INS alignment according to the manufacturer’s procedures and input flight plan waypoints and record. (The airplane’s present position needs to be loaded before the system can complete alignment. Do not move the airplane until the system is aligned and in the Navigate mode, if so equipped.) Present position accuracy should be known to within
0.1 NM.
(bb) Normal wind buffeting and movement of cargo and personnel will not affect alignment quality, but severe buffeting may lengthen the time required for alignment.
(cc) Continue operation in the NAV mode until a minimum of one hour has passed.
4 Pre-departure.
(aa) Recheck that other avionics operate normally and that all systems and instruments using INS information operate correctly.
(bb) Set up INS for a track from present position to any stored waypoint and check proper cockpit display unit (CDU) and instrument displays.
(cc) 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 warning condition is indicated. Record the time at which NAV mode is verified.
5 Departure Procedure.
224 (aa) Taxi. While taxiing the airplane, monitor INS performance with the CDU data selector set to display track angle and ground speed.
(bb) Takeoff. With the CDU set as required to observe ground speed, track angle, heading, and drift angle during takeoff and initial airborne conditions, verify normal and proper performance. Reasonable assessment may be accomplished using other instruments and avionics.
(cc) Initial track. Define a track from present position to first available waypoint and fly to intercept using instrument display or autopilot.
6 In Flight. During flight conditions, perform and/or check the following:
NOTE: Do not change system mode selector from “Navigate” during navigational tests. If a system is installed to provide attitude for the airplane system, the “ATT REF” mode should be checked after navigation mode tests are completed). When selector is changed to “ATT REF,” ascertain that all appropriate equipment annunciators indicate the disconnect.
(aa) 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 AFM.
(bb) If updating capabilities are provided, either as manual or automatic, use defined procedures to verify operation.
(cc) Operate the INS in both the manual and automatic modes for waypoint sequencing.
(dd) 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.).
(ee) Operate the CDU in all display settings and verify functions and operation.
(ff) Operate other avionics 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.
(gg) If basic accuracy has been previously established, a triangular course should be flown with a base leg length of at least one hour. The required accuracy is defined in paragraph 170c(9)(a), above.
7 Post Flight.
(aa) 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.
(bb) 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.
(cc) Turn system to off.
(e) Attitude (if provided to the basic instruments).
1 The system should be installed to provide pitch and roll data to ±1 degree relative to the airplane’s reference level.
2 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 that information as a limitation.
3 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 selftests 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.
(f) 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) Vertical Navigation (VNAV) System . A complete and comprehensive flight test evaluation is contained in AC 20-129, “Airworthiness Approval of Vertical Navigation (VNAV) Systems for use in the U.S. National Airspace System (NAS) and Alaska.”
dated September 12, 1988.
(11) Doppler Navigation System . Doppler Navigation System installed performance should be evaluated in accordance with AC 121-13 Change 2, “Self Contained Navigation Systems (Long Range),” dated December 21 1970. See also 14 CFR part 121, Appendix G, and paragraph 170c(9) of this AC, which addresses Inertial Navigation Systems.
(12) Area Navigation (RNAV) . Information pertinent to flight evaluation of area navigation is contained in AC 90-45A Change 2, “Approval of Area Navigation Systems for Use in the U.S. National Airspace System,” dated February 21, 1975. AC 90-100A, “US Terminal and En Route Area Navigation (RNAV) Operations” dated March 1, 2007; and AC 90-101, Approval Guidance for RNP Procedures with SAAAR” dated December 15, 2005. Note that the area navigation function is most often performed by the systems identified in paragraphs 170c(9), (10), (11), and (13) of this AC. Flight test requirements should be reviewed accordingly.
(13) Multi-Sensor Navigation System . A complete and comprehensive flight test evaluation is contained in AC 20-130A, “Airworthiness Approval of Navigation or Flight Management Systems Integrating Multiple Navigation Sensors,” dated June 14, 1995.
(14) Performance Management System (PMS) . A complete and comprehensive flight test evaluation is contained in AC 25-15 “Approval of Flight Management Systems in Transport Category Airplanes,” dated November 20, 1989.
(15) Flight Management System (FMS) . A complete and comprehensive flight test evaluation is contained in AC 25-15, “Approval of Flight Management Systems in Transport Category Airplanes,” dated November 20, 1989.
(16) Global Navigation Satellite System (GNSS) . A complete and comprehensive flight test evaluation is contained in AC 20-138A, “Airworthiness Approval of Global Navigation Satellite System (GNSS) Equipment,” dated December 22, 2003.
(17) Microwave Landing System (MLS) . [Reserved]
d. Procedures-Instruments and Displays .
(1) Flight and Navigation Instruments .
(a) Flight demonstration of the following pilots’ instruments should address evaluation with regard to display, function, and lighting:
1 Attitude Director Indicator (ADI)-primary and standby 2 Horizontal Situation Indicator (HSI) 3 Radio Magnetic Indicator (RMI) 4 Airspeed and Mach No. Indicator-primary and standby (if installed) 5 Altimeter-primary and standby 6 Instantaneous Vertical Speed Indicator (IVSI) 7 Standby Compass 8 Clock 9 Total Air Temperature Indicator (TAT) 1 0 Radio Altimeter (R/A)
(b) 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.
(c) In addition to the above instruments, the controls on the instrument, and those associated with changing the function of that instrument, should be evaluated.
(d) Day and night lighting of instruments should 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 AC 25-11A “Transport Category Airplane Electronic Display Systems,” dated June 21, 2007.
e. Procedures-Sensors and Warning Systems .
(1) Compass System .
(a) 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 43.13-1B Change 1, “Acceptable Methods, Techniques and Practices – Aircraft Inspection and Repair,” dated September 27, 2001, or latest revision.)
(b) Compare indicated heading information with known runway directions during takeoffs and landings.
(c) 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.
(d) 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) .
(a) 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 as discussed in this paragraph. Attitude and heading reference systems (AHRS) and inertial reference systems (IRS) are the types of systems envisioned.
(b) Validation of acceptable equipment installations includes, but is not limited to, the validation of proper installation considering the combined effects of temperature, altitude, electronic interference (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:
1 Vibration. Testing is generally accomplished on the sensor level to the criteria of RTCA DO-160F, “Environmental Conditions and Test Procedures for Airborne Equipment.” The structural mounting provisions on the airplane should provide assurance that the DO-160F 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.
2 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 should be addressed by analysis and demonstration, where applicable.
229 3 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 selftests 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.
(c) Multi-axis Failures - 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, fail-passive 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.
(d) 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 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.
(e) 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 should then be clear and unmistakable annunciation of the fact that this attitude condition exists.
(f) 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.
(g) Some additional flight evaluation suggestions follow:
1 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.
2 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.
3 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 en route, maneuvering, and coupled approach operations.
4 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.
5 Verify that the comparator monitor provides proper annunciation of attitude/heading disparity.
6 Verify that loss of a single power source does not cause loss of both attitude systems simultaneously.
7 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.
8 Verify that the system can be realigned in flight after being shutdown for more than 3 minutes (if applicable).
9 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.
1 0 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.
1 1 Verify that all controls, displays, and annunciations are satisfactorily identified, accessible, operable, and adequate for direct sunlight and night conditions.
NOTE: Some of the above items can be accomplished in a laboratory or simulator environment.
(h) Verify that there is no unacceptable mutual interference between the attitude system and other systems and equipment.
(3) Angle-of-attack (AOA) System .
(a) Pilot interface with this system w ill 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).
(b) Evaluate the displays during normal approaches and landings to assure reasonable/proper information is being presented for the particular flap setting and VREF speed over the gross weight and center of gravity range of the airplane (S/F should be centered).
(c) Conduct left and right sideslips that would be representative of normal operation on approach during crosswind landings or engine-out conditions.
(d) Determine that the malfunction indicators (flags) are appropriately displayed and are satisfactory annunciators.
(e) 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.
(f) Determine that instrument presentation is sufficiently damped to permit use in turbulent air and that hysteresis, if present, is acceptable.
(g) Verify that the AOA system neither contributes to, nor is affected by, radio frequency or electromagnetic interference (RFI/EMI).
(4) Air Data System .
(a) The air data system performance, with the exception of airspeed calibration, will generally be demonstrated qualitatively by observation of the output displays.
(b) Be particularly observant of barometric altimeter and vertical speed reversals as a function of rapid pitch attitude changes (e.g., takeoff rotation).
(c) Verify that the Mach/airspeed indicators operate smoothly throughout the complete speed envelope of the airplane.
(d) Observe correct operation of the overspeed warning indicator (barber pole) while approaching VMO/MMO speeds.
(e) Observe the static and total air temperature (SAT/TAT) indicators for reasonableness of data presentation.
(f) Verify that the air data system neither contributes to, nor is affected by RFI/EMI.
(5) Radio Altimeter System .
(a) 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.
(b) Under the measurement conditions described, verify that the altimeters:
1 Display altitude without loss of signal indications or excessive fluctuations, under the following conditions:
(aa) Pitch angle ±5 degrees about the mean approach attitude.
(bb) Roll angle zero to ±20 degrees.
(cc) Forward velocity from minimum approach speed up to 200 knots, in appropriate configurations.
(dd) At altitudes from 0 to 200 feet with sink rates of 0 to 15 feet/second, in landing, approach, and go-around configurations.
2 Track the actual altitude of the airplane over level ground without significant lag or oscillation.
(c) 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.
(d) 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.
(e) If a decision height (DH) functi on is provided, verify proper operation at minimum altitudes of 200 feet, 100 feet, and 50 feet.
(f) 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.
(g) 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.
(h) Verify that the radio altimeter system neither contributes to, nor is affected by RFI/EMI.
(6) On-Board Weight and Balance System .
(a) 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.
(b) 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.