f. 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.
g. 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 32.1.3.9.1 above.
7. Post Flight.
a. 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.
b. Pull INS primary power circuit br eaker and verify operation on battery for 5 minutes. Reset breaker and verify that the system functions normally and the waypoint data is correct.
c. Turn system to off.
32.1.3.9.5 Attitude (if Provided to the Basic Instruments).
1. The system should be installed to provide pitch and roll data to ±1° 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° 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 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 32-24
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.
32.1.3.9.6 EMC.
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.
32.1.3.10 Vertical Navigation (VNAV) System.
A complete and comprehensive flight test evaluation is contained in AC 20-138D, Change 2.
32.1.3.11 Area Navigation (RNAV).
Information pertinent to flight evaluation of RNAV is contained in AC 20-138D, Change 2. Operational approval guidance is contained in AC 90-100A, AC 90-101A, and AC 90-105A. The RNAV function is most often performed by the systems identified in paragraphs 32.1.3.9,
32.1.3.10, and 32.1.3.12 of this AC. Flight test requirements should be reviewed accordingly.
32.1.3.12 Multi-Sensor Navigation System.
A complete and comprehensive flight test evaluation is contained in AC 20-138D, Change 2.
32.1.3.13 Performance Management System.
A complete and comprehensive flight test evaluation is contained in AC 25-15.
32.1.3.14 Flight Management System.
A complete and comprehensive flight test evaluation is contained in AC 25-15.
32.1.3.15 Global Navigation Satellite System (GNSS).
A complete and comprehensive flight test evaluation is contained in AC 20-138D, Change 2.
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32.1.4 Procedures—Instruments and Displays.
32.1.4.1 Flight and Navigation Instruments.
32.1.4.1.1 Flight demonstration of the following pilots’ instruments should address evaluation with regard to display, function, and lighting:
Attitude director indicator (ADI)—primary and standby Horizontal situation indicator (HSI) Radio magnetic indicator (RMI) Airspeed and Mach number indicator—primary and standby (if installed) Altimeter-primary and standby Instantaneous vertical speed indicator (IVSI) Standby compass Clock Total air temperature (TAT) indicator Radio altimeter (R/A)
32.1.4.1.2 The instruments should be evaluated for suitability of location and performance during flight operation. Agreement between displayed functions, primary to standby, and the 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° and pitch angles from +25 to -10°.
32.1.4.1.3 In addition to the above instruments, the controls on the instrument, and those associated with changing the function of that instrument, should be evaluated.
32.1.4.1.4 Day and night lighting of instruments should be demonstrated. Evaluate for adequate illumination of the instruments and associated control panels and placards.
32.1.4.2 Electronic Display Systems.
A complete and comprehensive description and flight evaluation is contained in AC 25-11B.
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32.1.5 Procedures—Sensors and Warning Systems.
32.1.5.1 Compass System.
32.1.5.1.1 Conduct a ground compass swing and record deviations from correct magnetic headings at 15° 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, or latest revision.)
32.1.5.1.2 Compare indicated heading information with known runway directions during takeoffs and landings.
32.1.5.1.3 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.
32.1.5.1.4 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.
32.1.5.2 Attitude System—“Strapdown” (AHRS & IRS).
32.1.5.2.1 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.
32.1.5.2.2 Validation of acceptable equipment installations includes, but is not limited to, the validation of proper installation considering the combined effects of temperature, altitude, electromagnetic 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-160G, Environmental Conditions and Test 32-27
Procedures for Airborne Equipment. The structural mounting provisions on the airplane should provide assurance that the DO-160G 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.
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 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.
32.1.5.2.3 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 33.7 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 32-28
fail-passive cruise mode autopilots may not be truly fail-passive when reconfigured to reversionary modes.
32.1.5.2.4 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.
32.1.5.2.5 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.
32.1.5.2.6 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.
32.1.5.2.7 Some additional flight evaluation suggestions to follow:
1. While on the ground, determine that the attitude system provides usable attitude information through 360° of roll and pitch by rotating the platform through 360° 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.
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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 shut down 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.
10. 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 flag appear on the effected displays.
11. 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.
32.1.5.2.8 Verify that there is no unacceptable mutual interference between the attitude system and other systems and equipment.
32.1.5.3 Angle-of-attack (AOA) System.
32.1.5.3.1 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).
32.1.5.3.2 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 CG range of the airplane (S/F should be centered).
32.1.5.3.3 Conduct left and right sideslips that would be representative of normal operation on approach during crosswind landings or engine-out conditions.
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32.1.5.3.4 Determine that the malfunction indicators (flags) are appropriately displayed and are satisfactory annunciators.
32.1.5.3.5 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.
32.1.5.3.6 Determine that instrument presentation is sufficiently damped to permit use in turbulent air and that hysteresis, if present, is acceptable.
32.1.5.3.7 Verify that the AOA system neither contributes to, nor is affected by, radio frequency or electromagnetic interference (RFI/EMI).
32.1.5.4 Air Data System.
32.1.5.4.1 The air data system performance, except for airspeed calibration, will generally be demonstrated qualitatively by observation of the output displays.
32.1.5.4.2 Be particularly observant of barometric altimeter and vertical speed reversals as a function of rapid pitch attitude changes (e.g., takeoff rotation).
32.1.5.4.3 Verify that the Mach/airspeed indicators operate smoothly throughout the complete speed envelope of the airplane.
32.1.5.4.4 Observe correct operation of the overspeed warning indicator (barber pole) while approaching VMO/MMO speeds.
32.1.5.4.5 Observe the static and total air temperature (SAT/TAT) indicators for reasonableness of data presentation.
32.1.5.4.6 Verify that the air data system neither contributes to, nor is affected by RFI/EMI.
32.1.5.5 Radio Altimeter System.
32.1.5.5.1 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.
32.1.5.5.2 Verify that the altimeters display altitude without loss of signal indications or excessive fluctuations, under the following measurement conditions:
Pitch angle ±5° about the mean approach attitude.
Roll angle zero to ±20°.
Forward velocity from minimum approach speed up to 200 knots, in appropriate configurations.
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At altitudes from 0 to 200 feet with sink rates of 0 to 15 feet per second, in landing, approach, and go-around configurations.
32.1.5.5.3 Under the measurement conditions described above, verify that the altimeters track the actual altitude of the airplane over level ground without significant lag or oscillation.
32.1.5.5.4 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.
32.1.5.5.5 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.
32.1.5.5.6 If a decision height (DH) function is provided, verify proper operation at minimum altitudes of 200 feet, 100 feet, and 50 feet.
32.1.5.5.7 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.
32.1.5.5.8 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.
32.1.5.5.9 Verify that the radio altimeter system neither contributes to, nor is affected by RFI/EMI.
32.1.5.6 Onboard Weight and Balance System.
AC 20-161 provides guidance for certification of onboard weight and balance systems dated April 11, 2008.
32.1.5.7 Central Aural Warning System (CAWS).
32.1.5.7.1 Most CAWS functions include aural (and voice, in some installations) warnings of the typical airplane conditions specified in table 32-4 below:
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Table 32-4. Types of Aural Warnings
Airplane Condition Aural Warning Engine Fire Altitude Advisory APU Fire Speedbrake Overspeed Autopilot Disconnect Takeoff Flap/Slat Overspeed Stall Cabin Altitude Landing Gear Stabilizer-In-Motion Evacuation Autothrottle Disconnect
32.1.5.7.2 Pilot call and SELCAL tones may also be included, although they are advisories as opposed to warnings.
32.1.5.7.3 The logic for this system usually resides in another computer (e.g., the flight warning computer or master caution/master warning computer) in which case the two systems should be evaluated simultaneously. However, if the logic and/or computer governing activation of the warnings, priority logic, inhibits, etc., are contained in the CAWS itself, refer to the section covering the computers noted above for additional guidance on flight evaluation. Compliance may be shown by equivalent tests in a flight simulator or by bench tests.
32.1.5.7.4 Individual activation of each warning tone should be accomplished in flight under the most adverse noise conditions appropriate to the warning
(e.g., VMO/MMO for overspeed, high speed descent for APU fire, acceleration following takeoff for flap/slat overspeed, etc.). Test switches and/or the self-test feature of the system would be an appropriate means to activate the warning.
32.1.5.7.5 The flight evaluation will be a pilot qualitative assessment of the clarity, tone, and volume of each warning.
32.1.5.8 Overspeed Warning System.
This system should be demonstrated in conjunction with the airspeed system calibration and should exhibit that it is consistently within the prescribed tolerances. The system tolerance is from VMO to VMO + 6 knots in the airspeed limited range, and from MMO to MMO + 0.01 M in the Mach limited range.
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32.1.5.9 Altitude Advisory (or Altitude Alerting) System.
32.1.5.9.1 A ground test should be conducted to check the adequacy of the required pre-flight procedure provided in the AFM.
32.1.5.9.2 A flight demonstration should be conducted at low, mid, and high altitudes to verify performance of the intended function of:
Alerting to, and Alerting to an uncommanded departure from an assigned (selected) altitude.
32.1.5.9.3 Low and high rates of vertical speed should be included in the evaluation with captures from above and below the pre-selected altitude.
32.1.5.9.4 Determine adequacy of the alert light location and its visibility under various lighting conditions, and adequacy of the aural warning.
32.1.5.9.5 A flight simulator may be used to perform appropriate tests.
32.1.5.10 Terrain Awareness and Warning System.
AC 25-23 contains guidance for flight test considerations for TAWS approvals.
32.1.5.11 Master Caution System (MCS)/Master Warning System (MWS).
32.1.5.11.1 This system provides inputs to master warning and master caution lights located on the glareshield in front of each pilot and to the CAWS. (See paragraph 32.1.5.8 of this AC.) It contains the sensors, switches and logic required to assess the requirement for the warning and caution annunciations.
32.1.5.11.2 The system may contain logic to assess priority of the various warnings in order to sequence simultaneous actuation. If so, ground, flight, or bench tests may have to be conducted to verify this logic.
32.1.5.11.3 Some cautions may be inhibited during various phases of flight, particularly during the final stages of a Category III automatic landing operation (below alert height). If so, ground or flight tests may have to be structured to verify the inhibit logic. Also, certain warnings may be inhibited during the high-speed stages of a takeoff roll. If this feature is incorporated, it too should be carefully substantiated.
32.1.5.11.4 Simulate failure of selected sensor inputs to verify that the MCS/MWS computer detects the faults and illuminates the fail annunciator (if installed).
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32.1.5.11.5 Verify that there is no unacceptable mutual interference between the MCS/MWS and other systems and equipment.
32.1.5.12 Flight Warning Computer (FWC).
32.1.5.12.1 The FWC is a key subset of the overall warning, caution, and alerting system on the more modern airplanes. It is a microprocessor-based system that works in conjunction with, or totally replaces, the CAWS and MCS/MWS.
32.1.5.12.2 The FWC incorporates sufficient computer capacity to perform many more functions than the MCS/MWS and, in addition to outputting to the visual warning displays and aural warning speakers, outputs hundreds of alerts/messages to other display media, most often an electronic display (multi-function CRT on the instrument panel or the CDU portion of the FMS).
32.1.5.12.3 The added computer capacity also enables incorporating rather complex inhibit algorithms which, in turn, dictates a requirement for structuring a very complex ground test to verify operation of the inhibits. Review of the system description document will probably be required before any ground, flight, or bench tests can be defined.
32.1.5.12.4 The flight evaluation should incorporate the tests described under CAWS and MCS/MWS.
32.1.5.12.5 Verify that the self-test feature functions properly.
32.1.5.12.6 Simulate failure of selected sensor inputs to verify that the FWC detects the faults and illuminates the FWC fail annunciator.
32.1.5.12.7 Demonstrate freedom from nuisance warnings by observing and logging nuisance events as they occur throughout the flight demonstration program. Assess the status at the end of the program by examining the log.
32.1.5.13 Stall Warning Computer (SWC).
32.1.5.13.1 The SWC should be evaluated in conjunction with the stall tests described in chapter 8 of this AC. Adequacy of the various warnings (stick shaker, warning light, aural (and voice) warning) should be assessed, in addition to the speed schedules.
32.1.5.13.2 Demonstrate freedom from nuisance warnings during the takeoff, landing, and maneuvering flight tests described in chapter 8 of this AC. Further testing may be required to show freedom from nuisance warnings by the test requirements in chapter 42 of this AC, if special airworthiness approvals are sought.
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