(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.
(8) Central Aural Warning System (CAWS) .
(i) Most CAWS functions include aural (and voice, in some installations) warnings of the following typical airplane conditions:
Engine Fire Altitude Advisory APU Fire Speedbrake Overspeed Autopilot Disconnect Takeoff Flap/Slat Oversp eed Stall Cabin Altitude Landing Gear Stabilizer-In-Motion Evacuation Pilot Call and SELCAL tones may also be included, although they are "Advisories" as opposed to "Warnings."
(ii) The logic for this system usually resides in anot her 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.
(iii) Individual activation of each warning tone should be accomplished in flight under the most adverse noise conditions appropriate to the warning (e.g., V MO / 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.
(iv) The flight evaluation will be a pilot qualitative assessment of the clarity, tone, and volume of each warning.
(9) 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 M MO to MMO + 0.01M in the Mach limited range.
(10) Altitude Advisory (or Altitude Alerting) System .
(i) A ground test should be conducted to check the adequacy of the required pre-flight procedure provided in the Airplane Flight Manual.
(ii) A flight demonstration should be conducted at low, mid, and high altitudes to verify performance of the intended function of:
(A) Alerting to an impending capture of a pre-selected altitude; and (B) Alerting to an uncommanded departure from an assigned (selected) altitude.
(iii) Low and high rates of vertical speed should be included in the evaluation with captures from above and below the pre-selected altitude.
(iv) Determine adequacy of the alert light location and its visibility under various lighting conditions, and adequacy of the aural warning.
(v) A flight simulator may be used to perform appropriate tests.
(11) Ground Proximity Warning System (GPWS) .
(i) Initial approvals of new GPWS should include flight evalua tion with an instrumented airplane.
The instrumented flight test of a computer type may be demonstrated on one airplane and the results made applicable to any other airplane with only a functional test being required.
(ii) Initial approval of new indi vidual Technical Standard Order approved ( TSO'd) sensors, which have an equivalent function and interface as previously approved sensors, and are interfaced with a previously approved computer, may be approved by ground test. Applicable sensors (systems) are the air data computer system (ADC), the radio altimeter system (R/A), and the Instrument Landing System (ILS).
(iii) Subsequent installations of identical computers and interfacing sensors in other airplane types will require a ground test and an uninstrumented functional flight test to determine proper functioning, freedom from interference, adequate sound and light levels, inhibit switch details, etc.
(iv) The functional flight test must include construction of a profile to selectively actuate each mode that the applicant elects to certify (not all modes are required by the FAA). Refer to the equipment supplier's system description to aid in constructing the flight profile.
(v) Evaluate the GPWS warning light and other annunciators for visibility in day and night lighting conditions.
(vi) Evaluate the sound level of the aural GPWS warning at V MO with headsets on and with and without simultaneous radio communications.
(vii) Verify that the self-test function and the flap override swi tch perform their intended functions.
The gear override switch should be verified as well.
(viii) Simulate failures of the ADC, R/A's and ILS systems (open circuit breakers) and verify that the GPWS detects the faults and illuminates the GPWS fail annunciator.
(ix) Verify that conducting a radio altimeter self-test in flight will not induce a false GPWS failure warning.
(x) Verify that the GPWS does not adversely interfere or interact with other airplane systems.
(xi) Verify that the GPWS is free of inadvertent activation during electrical bus or generator switching.
(xii) Verify that the caution & warning priorities are assessed and that they include systems such as windshear. Example, windshear has priority over TCAS.
(12) Enhanced Ground Proximity Warning System (EGPWS). (RESERVED.)
(13) Master Caution/Master Warning System (MCS/MWS) .
(i) This system provides inputs to master warning (MW) and master caution (MC) lights located on the glareshield in front of each pilot and to the CAWS (see paragraph 170 e(8)). It contains the sensors, switches and logic required to assess the requirement for the warning and caution annunciations.
(ii) 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.
(iii) 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.
(iv) Simulate failure of selected sensor inputs to verify that the MCS/MWS computer detects the faults and illuminates the fail annunciator (if installed).
(v) Verify that there is no unacceptable mutual int erference between the MCS/MWS and other systems and equipment.
(14) Flight Warning Computer (FWC) .
(i) 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 MC/MW systems.
(ii) 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 Flight Management System).
(iii) The added computer capacity also en ables 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.
(iv) The flight evaluation should incorporate the tests described under CAWS and MC/MW systems.
(v) Verify that the self-test feature functions properly.
(vi) Simulate failure of selected sensor inp uts to verify that the FWC detects the faults and illuminates the FWC fail annunciator.
(vii) 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.
(15) Stall Warning Computer (SWC) .
(i) The SWC should be evaluated in conjunction with the stall tests described in Chapter 2, Section 6 of this Advisory Circular. Adequacy of the various warnings (stick shaker, warning light, aural (and voice) warning) should be assessed, in addition to the speed schedules.
(ii) Demonstrate freedom from nuisance warnings during the takeoff, landing, and maneuvering flight tests described in Chapter 2 of this Advisory Circular. Further testing may be required to show freedom from nuisance warnings by the test requirements in Chapter 8 of this AC if special airworthiness approvals are sought.
(16) Takeoff Warning System (TOW) .
(i) The TOW system is usu ally a subset of the CAWS or MCS/MWS or FWC and is generally inhibited at nose wheel liftoff during the takeoff maneuver. It is usually armed when the squat switch indicates that the airplane is on the ground and one or more power levers are advanced beyond a prescribed position or engine thrust setting. The system alarms if the flaps/slats/stabilizer/spoiler/brakes are not set correctly for takeoff.
(ii) Determine that each event is annunciated properly by conducting a static ground test.
(iii) Flight evaluation during takeoff roll should be a qualitative assessment of the clarity, tone, and volume of each warning. Some systems incorporate a TEST switch, which facilitates activation during takeoff. If this feature is not available, it may suffice to check the audio level of a warning during touch and go landings (e.g., the "stabilizer" warning while accelerating to V R speed).
(iv) Simulate failure of selected sensor inputs to verify that the TOW system detects the faults and illuminates the fail annunciator (if installed).
(v) Verify that the system "arm status" is not affected by electrical power transients, which occur when switching from ground or APU power, during engine start, following bus priority checks, etc.
(vi) Verify that there is no unacceptable mutual interference between the TOW system and other airplane systems and equipment.
(17) Instrument Comparator System .
(i) If an instrument comparator system is installed and activated, ensure the system monitors validity of and compares (side 1 vs. side 2) pitch and roll attitude, heading, radio altitude, localizer, and glideslope deviation information used for display in the cockpit instruments.
(ii) When these parameters exceed a preset threshold or trip level, or if an invalid condition is detected in any of the displayed information, an appropriate annunciator to that parameter will be illuminated.
(iii) A ground test on the airplane should be conducted to demonstrate performance of intended functions and validate the detector threshold levels. If annunciation is suppressed under certain conditions, these conditions should also be checked.
(iv) A flight evaluation should also be conducted to determine adequacy of the annunciator light locations and visibility under various lighting conditions and that the annunciators are properly identified.
(v) During flight, verify that the system is free from nuisance warnings and is compatible with other electronic systems.
(18) Reactive Windshear Warning System . A complete and comprehensive demonstration program (simulator and flight test) is contained in Advisory Circular 25-12, "Airworthiness Criteria for the Approval of Airborne windshear Warning Systems In Transport Category Airplanes," dated November 2, 1987.
(19) Traffic Alert and Collision Avoidance System . A complete and comprehensive demonstration program (laboratory/ground/flight test) is contained in Advisory Circular 20-131A, "Airworthiness And Operational Approval of TCAS II And Mode S Transponders," dated March 29, 1993.
f. Procedures-Recording Systems .
(1) Cockpit Voice Recorder (CVR) .
(i) Demonstrate operation of the CVR self-test function and bulk erase feature.
(ii) During ground operations prior to and after engine start; du ring takeoff, climbout, cruise at VMO / MMO, and during landing approach. Except where specifically noted, obtain the following CVR recordings:
(A) Flightcrew conversations using the area microphone.
(B) Flightcrew conversations using the oxygen mask microphones and the boommicrophone/handset, and the PTT hand-held microphones in flight and during ground operations.
(C) Radio transmissions by each crew member during cruise.
(D) Audio signals identifying navigation aids (through cockpit speakers and on radio channels) during landing approach.
(E) PA announcement during cruise.
(F) Selective ( based on analysis) aural warning signals (see CAWS) during appropriate flight phases .
(iii) Use CVR circuit breaker to limit CVR re cording to available capacity .
(iv) Verify that all controls, displays, and annunciators consistent with normal and abnormal crew procedures, are satisfactorily identified, accessible, operable, and visible.
(v) Electro-Magnetic Compatibility (EM C). Verify by observation that no adverse effects are present in the required flight systems.
(vi) Following the flight, test the CVR audio recording in an appropriate laboratory, and evaluate for clarity of the recorded messages. Preferably the evaluators should not have advance information concerning the flightcrew conversations.
(vii) Verify that the hot mike feature at each position allows for recording regardless of the position of the interphone transmit key switch (§ 25.1457(c)(5).
(2) Digital Flight Data Recorder (DFDR) .
(i) Demonstrate operation of the DFDR self-test feature and, prior to collection of test data, set the trip number and date on the code switches of the Trip and Data Encoder, if so equipped.
(ii) The certification data will be derived automatically during any appropriate flight that encompasses operation throughout the full flight envelope. Hand recorded data of heading (magnetic or true), barometric altitude, pitch attitude, indicated airspeed, and time (UTC) should be collected at several stabilized flight conditions for subsequent data correlation purposes.
(iii) Following the flight, the data should be retrieved for correlation and accuracy verification.
(iv) Demonstrate performance of intended func tion of the DFDR for all parameters specified in the operating rules (refer to Appendix B of part 121 for the relevant data ranges, accuracies, and recording intervals.)
Accuracy will be verified at stabilized points by correlating against, and comparing with, an approved instrumentation system or manual notes.
(v) Electro-Magnetic Compatibility (EMC). Verify by observation that no adverse effects are present in the required flight systems.
g. Engine Interfacing Systems . [RESERVED]
h. Stability Augmentation Systems . [RESERVED]
i. All Weather Operation (Reduced Visibility) Systems . [RESERVED]
171. FLIGHT AND NAVIGATION INSTRUMENTS - § 25.1303
a. Explanation.
(1) Section 25.1303(b)(1) requires an airspeed indicator to be visible at each pi lot station. Additionally, if airspeed limitations vary with altitude, airspeed indicators must have a maximum allowable airspeed indicator showing the variation of V MO with altitude. In practice, this maximum allowable airspeed indicator requirement is rarely complied with in a literal sense. The variation in V MO is not normally depicted literally as a function of altitude, but rather as a radial marker on the airspeed indicator whose position varies as a function of altitude. This presentation of the maximum allowable airspeed limit, as a function of altitude, is approved based on a substantiation of equivalent safety with the literal interpretation of the rule.
(2) Production tolerances for speed warning devices at V MO/MMO are required in accordance with §
25.1303(c) and SR450A.
(3) Section 21.127(b)(2) requires each production flight test to include "An operational check of each part or system operated by the crew while in flight to establish that during flight, instrument readings are within normal range." Nowhere in these requirements is there any inference that the finite performance or quantitative limits, defined during type certification, need to be determined for each production airplane.