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.
(4) SR-450A 1(b) and § 25.1303(c) require that turbine-powered airplanes be equipped with a speed warning device that will provide aural warning whenever the speed exceeds V MO plus 6 knots or M MO + 0.01M. The regulations specify that the upper limit of the production tolerances permitted for the warning device shall be at a speed not greater than the prescribed warning speed.
b. Procedures . The applicant should substantiate, by appropriate ground and/or flight tests with possible production instrument error corrections, that the system operates within the boundaries established by the regulation.
Understanding that other procedures may be acceptable, this could be accomplished in accordance with the following:
(1) If the maximum allowable airspeed varies as a function of altitude, and is indicate d by means other than a maximum airspeed versus altitude presentation in the airspeed indicator, it should be substantiated that the chosen means of maximum allowable airspeed indication provides an accurate presentation of the relationship between VMO and altitude. The indicated V MO should not exceed the actual V MO for any altitude by more than six knots. This is particularly important for airplanes that have a step function in V MO at an altitude of 10,000 feet due to bird strike considerations below that altitude.
(2) During production flights, the speed warning device shall provide warning at a speed equal to or less than VMO + 6 knots or M MO + 0.01M, minus the sum of the values stipulated in ( i) and (ii) below:
(i) The possible error of the airs peed indicator or Mach meter used as the reference.
(ii) Three knots or 0.005M, if the airspeed indicator or Mach meter used as the reference has a static or pitot pressure source different from the pressure sources of the warning devices.
(3) The speed warning device shall be shown to comply with the requirements of ( i) and (ii) below:
(i) During ground test using calibrated reference instruments, the installed speed warning device shall provide warning at a speed equal to or less than V MO + 6 knots or M MO + 0.01M. The procedure for this test should be processed in accordance with § 21.127(b)(5) or § 21.143(a)(3).
(ii) For each type design, a test should be conducted to show satisfactory correlation between the operation of the warning device during flight and the data obtained during the ground tests.
172. POWERPLANT INSTRUMENTS - § 25.1305 . [RESERVED]
173. MISCELLANEOUS EQUIPMENT - § 25.1307 . [RESERVED]
174. EQUIPMENT, SYSTEMS, AND INSTALLATIONS - § 25.1309 .
a. Explanation. The following procedures outline and paraphrase the appropriate provisions of § 25.1309.
Further definition and explanation, if required, may be found in Part 25 of the Federal Aviation Regulations and in Advisory Circular 25.1309-1A, "System Design and Analysis," dated June 21, 1988.
b. Procedures .
(1) Evaluate functioning of required installed equipment to verify that performance is as intended under any foreseeable operating and environmental conditions.
(2) Evaluate failure conditions, as appropriate, to determine their impact on the capability of the airplane or the ability of the crew to operate it.
(3) Review, as appropriate, any design analyses, proposals, studies, or tests that correlate probabilities of failure condition occurrence with the effects of those failure conditions, to determine that they are properly categorized for the appropriate criticality level.
(4) Verify that adequate warnings are provided of unsafe conditions, and that these warnings enable the flightcrew to take appropriate corrective action with a minimum of error.
(5) For probable operating combinations of required electrical installations, verify that power is provided for probable durations:
(i) Under normal conditions;
(ii) After failure of any one generato r, converter, or storage device;
(iii) After failure of one engine on a two-engine airplane; and
(iv) after failure of two engines on airplanes with three or more engines.
(6) For probable operating combinations of required electrical installati ons, which must be provided with an alternate source of power, verify that power is provided for probable durations after failure of any one power system.
Section 2. INSTRUMENTS: INSTALLATION
175. ARRANGEMENT AND VISIBILITY - § 25.1321. [RESERVED]
176. WARNING, CAUTION, AND ADVISORY LIGHTS - § 25.1322. [RESERVED]
177. AIRSPEED INDICATING SYSTEM - § 25.1323.
a. Explanation.
(1) Methods. Unless a calibrated reference system is provided, the airspeed system should be calibrated throughout as wide a range as necessary to cover the intended flight tests.
The procedures of this section are for the purpose of showing compliance with § 25.1323(b) and are not intended to cover the speed range of the flight tests. If an alternate airspeed indicating system is provided, it should be calibrated. The airspeed indicating system should be calibrated in accordance with the following methods:
(i) The tests should be conducted in stabilized flight at airspeeds throughout the speed range for the airplane configurations to be tested. The airplane's airspeed system should be calibrated against a reference airspeed system or a groundspeed course.
(ii) A reference airspeed system should consist of either of the following:
(A) An airspeed impact pressure and static pressure measurement device (or devices) that are free from error due to airplane angular changes relative to the direction of the free stream or due to slipstream variation resulting from changes in airplane configuration or power. In addition, the device or devices should have a known calibration error when located in the free stream; or (B) Any other acceptable airspeed calibration method (e.g., the altimeter method of airspeed calibration).
(iii) When establishing the airplane's true airspeed by means of the groundspeed course, flight between the two reference points should be made at constant airspeed in two successive runs in opposite directions to eliminate the effect of wind. The runs should be made only in stable wind. The time to make the runs should be obtained by means of some calibrated device. The speed runs should not be made nearer the ground surface than a wing span's length.
(iv) If an alternate system is provided, it may be calibrated against either the reference system or the airplane's system.
(v) Airspeed Lag. With the advent of electronic instruments in the cockpit, the pneumatic signals from the pitot and static sources are processed and digitized in the Air Data Computer (ADC) and then filtered and transported to the cockpit display. As a result of the data processing and filtering, the associated time lag, and, consequently, airspeed lag at the cockpit display, can be an important consideration in the airspeed indicating system calibration during ground acceleration. As stated in § 25.1323(b), the calibration for an accelerated takeoff ground
run must determine the "system error" which is the relation between indicated and calibrated airspeeds. The system error is the sum of the pneumatic lag in the pressure lines, airspeed lag due to time lags in processing the data, and static source, position error.
(A) Airspeed lag must be measured during ground acceleration tests or determined by analysis. Increments should be developed for a range of airplane gross weights considering airspeed lag at V1 and the associated increase in accelerate-stop and takeoff distances due to lag. The error due to lag in the airspeed indicating system during ground acceleration should not be greater than 3.0 knots throughout the takeoff operating envelope of the airplane. Furthermore, an increase in the takeoff distance or the accelerate-stop distance as a result of airspeed lag should not exceed 100 ft. The 3 knots limitation is intended to establish the maximum acceptable systematic error. Even though the lag may be within the 3 knots limit, an airspeed correction may be required to stay within the 100 ft. of increased distance.
(B) Corrections may be applied directly in the ADC or they may be introduced via the ground airspeed calibration provided in the Airplane Flight Manual (AFM). If corrections are applied directly in the ADC, it is possible to display calibrated airspeed in the cockpit. Furthermore, if acceleration data are input, the airspeed error can be computed and accounted for in real time, assuming the time lag is known. The alternative would be to use an airspeed lag increment derived from calibration tests that would represent a range of conditions within the takeoff envelope. After correction, an increase in distance due to lag should be less than 100 ft throughout the takeoff envelope, whether applied in the ADC or AFM.
Consideration should be given to short field, lighter weight takeoffs (higher acceleration), as well as maximum weight and higher V1 speeds, in deriving the increment.
(2) Configuration. Airspeed calibration tests should be conducted in the following configurations:
(i) Weight - between maximum takeoff and maximum landing.
(ii) C. G. position - optional.
(iii) Takeoff configuration(s) - ground roll.
(iv) Wing flaps and landing gear - all combinations of positions used to show compliance with the takeoff, climb, and landing requirements of 14 CFR part 25.
(v) Thrust - as required.
b. Procedures.
(1) Any one or any desired combination of the procedures in subparagraphs (2) through
(4) of this paragraph may be used for calibrating the airspeed indicating system. The airspeed should be measured or determined simultaneously from the airplane's system and the reference system during stabilized runs for at least five speeds spaced throughout the speed range, the lowest not to exceed 1.3 VS. The highest speed should not exceed VMO/MMO. The speed spread between the test speeds should be limited to 10 knots from VS to 1.6 VS or placard speed, and 20 knots from 1.6 VS to VMO.