(4) Section 25.1303(c) requires that turbine-powered airplanes be equipped with a speed warning device that will provide aural warning whenever the speed exceeds VMO plus 6 knots or MMO + 0.01 M. The regulations specify that the upper limit of the production tolerances permitted for the warning device must 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 indicated 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 VMO should not exceed the actual VMO for any altitude by more than six knots. This is particularly important for airplanes that have a step function in VMO at an altitude of 10,000 feet due to bird strike considerations below that altitude.
(2) During production flights, the speed warning device must provide warning at a speed equal to or less than VMO + 6 knots or MMO + 0.01 M, minus the sum of the values stipulated in (a) and (b) below:
(a) The possible error of the airspeed indicator or Mach meter used as the reference.
(b) Three knots or 0.005 M, 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 must be shown to comply with the requirements of § 25.1303(c), as explained in paragraph (a) and (b) below:
(a) During ground test using calibrated reference instruments, the installed speed warning device must provide warning at a speed equal to or less than VMO + 6 knots or MMO + 0.01M. The procedure for this test should be processed in accordance with § 21.127(b)(5) or § 21.143(a)(3).
(b) 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 and in AC 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 an alyses, 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:
(a) Under normal conditions;
(b) After failure of any one generator, converter, or storage device;
(c) After failure of one engine on a two-engine airplane; and
(d) After failure of two engines on airplanes with three or more engines.
(6) For probable operating combinations of required electrical installations that must be provided with an alternate source of power in accordance with § 25.1331(a), 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:
(a) 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.
(b) A reference airspeed system should consist of either of the following:
1 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
2 Any other acceptable airspeed calibration method (e.g., the altimeter method of airspeed calibration).
(c) 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 wingspan’s length.
(d) If an alternate system is provided, it may be calibrated against either the reference system or the airplane’s system.
(e) An acceptable means of compliance when demonstrating a perceptible speed change between 1.23 VSR to stall warning speed (§ 23.1323(d)) is for the rate of change of IAS with CAS to be not less than 0.75.
(f) An acceptable means of compliance when demonstrating a perceptible speed change between VMO to VMO + 2/3 (VDF - VMO) (§ 23.1323(e)) is for the rate of change of IAS with CAS to be not less than 0.50.
(g) 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
244 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.
1 Airspeed lag should 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.
2 Corrections may be applied directly in the ADC or they may be introduced via the ground airspeed calibration provided in the 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:
(a) Weight - between maximum takeoff and maximum landing.
(b) C. G. position - optional.
(c) Takeoff configuration(s) - ground roll.
(d) Wing flaps and landing gear - all combinations of positions used to show compliance with the takeoff, climb, and landing requirements of part 25.
(e) 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 be less than 1.23 VSR. The highest speed should not exceed VMO/MMO.
245 The speed spread between the test speeds should be limited to 10 knots from 1.23 VSR to 1.5 VSR or VFE, and 20 knots from 1.5 VSR to VMO.
(2) Speed course : The airspeed, power, and altitude should be stabilized before entering the speed course. Constant airspeed should be maintained during each run. The runs should be made in both directions on reciprocal headings for each speed over the speed course.
The following data should be recorded:
(a) Time of day at beginning of run.
(b) Time to make run.
(c) Pressure altitude.
(d) Ambient air temperature.
(e) Airspeed at several intervals during run.
(f) Wing flap position.
(g) Landing gear position.
(h) Course distance.
(3) Reference airspeed system : Stabilized runs at the test speeds listed in this paragraph should be made. The airspeed from the airplane’s airspeed system and the reference airspeed system should be read simultaneously. The following data should be recorded:
(a) Time of day.
(b) Airplane’s indicated airspeed.
(c) Reference indicated airspeed.
(d) Pressure altitude.
(e) Ambient air temperature.
(f) Wing flap position.
(g) Landing gear position.
(4) Other acceptable airspeed calibration methods . Stabilized flight runs at the test speeds should be made, and the necessary data recorded, to establish the airplane’s airspeed system error and the configuration of the airplane. Calibration methods may also include airspeed boom, static trailing cone, and radar range.
(5) The procedures presented in this paragraph pertain to the calibration of the airspeed indicating system during takeoff ground acceleration. In particular, airplanes with electronic instruments in the cockpit must account for the airspeed lag at the cockpit display associated with data processing and filtering as required by § 25.1323(g). The airspeed indicating system should not have a lag in excess of 3 knots at the V1 speed during any takeoff condition.
Furthermore, if airspeed lag causes an increase of more than 100 ft. in takeoff or accelerate-stop distances, a lag correction should be applied to the airspeed indicating system. Airspeed lag should be determined by one of the following methods:
(a) Conduct ground acceleration tests for a range of airplane gross weights to calibrate IAS at the cockpit display against the reference CAS. Determine airspeed lag from the calibration data by comparing the cockpit displayed airspeed with the reference calibration speed for a given gross weight and V1 speed.
(b) Determine airspeed lag by analysis using a computer program suitable for AFM development. Compute takeoffs for a range of gross weights to determine the acceleration at V1. Calculate airspeed lag at V1 for a corresponding acceleration and a known time lag due to data processing and filtering. The analysis should also consider other sources of airspeed lag as appropriate, such as the pneumatic lag in the pressure lines for the pitot and static sources.
(6) Having established the calibration data, one acceptable method of adjusting for airspeed lag is to apply corrections directly in the ADC data processing to result in a lagcorrected airspeed at the cockpit display. Another would be to include an airspeed lag correction in the takeoff ground speed calibration of IAS vs. CAS in the AFM. A single airspeed lag increment can be developed as the correction for the range of gross weights and corresponding accelerations at V1. This increment, when applied to the calibration, should result in no more than a 100 ft. increase in takeoff or accelerate-stop distances due to airspeed lag for any takeoff condition. A more accurate correction would result from presenting airspeed lag as a function of airplane acceleration based on the calibration data. If acceleration data are available in the ADC, a real time correction for lag during the takeoff can be applied in the data processing.
178. Static Pressure Systems - § 25.1325(d) and (e).
a. Explanation .
(1) If the altimeter installation is of the pressure type, its operation will be affected by any error that exists in the measurement of the static air pressure. Since the accuracy of the altimeter is of utmost importance, the static air vent system should be calibrated. If separate or alternate vent systems are employed for the altimeter and airspeed indicator, separate calibrations are required. Where the altimeter, rate of climb indicator, and airspeed indicators are vented to the same static systems, the altimeter calibration may be made in conjunction with the airspeed calibrations.
(2) The theoretical relationship between airspeed error and altimeter error may be used to derive an altimeter calibration from the airspeed calibration if both use the same static vent,
247 provided the total head sensing installation is designed to provide true readings over the angleof-attack range involved.
b. Procedures . None.
179. Pitot Heat Indication Systems - § 25.1326. [Reserved]
180. Magnetic Direction Indicator - § 25.1327. [Reserved]
181. Automatic Pilot System - § 25.1329.
a. Explanation .
(1) On most modern airplane installations, the autopilot, automatic thrust control (if provided), and flight director are integrated and are identified as a flight guidance system (FGS).
Although a HUD system, if installed, is generally a separate system, evaluation of its modes of operation would be identical to a flight director system. For these reasons, the systems flight test evaluation criteria provided here also apply to HUDs and head up guidance systems (HGS).
(2) The system pertains to equipment or combinations of equipment that provide the means for airplane control and guidance, whether by means of pilot supervised automatic systems or through guidance to the pilot for manual control. This includes sensors, computers, power supplies, servomotors/actuators, and associated wiring.
(3) The system should include any indications and controllers necessary for the pilot to manage and supervise the system.
(4) Additional guidance material related to FGS operation, testing, and approval is contained in the following FAA Advisory Circulars:
(a) AC 25-15, “Approval of Flight Management Systems in Transport Category Airplanes,” dated November 20, 1989.
(b) AC 25.1329-1B, “Approval of Flight Guidance Systems,” dated July 17, 2006.
(c) AC 91-16, “Category II Operations - General Aviation Airplanes,” dated August 7, 1967.
(d) AC 120-28D, “Criteria for Approval of Category III Weather Minima for Takeoff, Landing, and Rollout,” dated July 13, 1999.
(e) AC 120-29A, “Criteria for Approval of Category I and Category II Weather Minima for Approach,” dated August 12, 2002.
b. Procedures .
(1) General .
(a) The evaluation of an FGS should primarily concentrate on system intended function and safe operation. In certain areas, degree of performance will require specific criteria commensurate with the type of operation. Criteria and requirements for flight should consider the applicant’s engineering analysis. The combination of analysis and flight evaluation will form the whole of the certification requirement and, as such, should be in harmony and provide full evaluation. The flight evaluations supplement the analysis as required in both the representative operating conditions of the airplane and for critical operating conditions. The critical conditions are determined from analysis of the total FGS/airplane modes of operation and configuration conditions.
(b) The requirement for demonstrating safe operation normally should include induced failures during flight. The requirement for failure demonstrations is also an outgrowth of the analysis submitted by the applicant and is a result of the particular design being evaluated
(e.g., consideration should be made for multiple channel systems specifically designed to be selfadaptive to failure conditions).
(c) The automatic pilot system should be installed and adjusted so that the system tolerances established during certification tests can be maintained in normal operation. This may be assured by conducting flight tests at the extremes of the tolerances. Those tests conducted to determine that the automatic pilot system will adequately control the airplane should establish the lower limit, and those tests to determine that the automatic pilot will not impose dangerous loads or deviations from the flight path should be conducted at the upper limit. Appropriate airplane loading to produce the critical results should be used.
(d) Means should be provided to indicate to the flightcrew the current mode of operation and any modes armed by the pilot. Selector switch position is not acceptable as a means of indication.
(e) The system should be demonstrated to perform its intended function in all configurations in which it may be used throughout all appropriate maneuvers and environmental conditions, including turbulence, unless an appropriate operating limitation or statement is included in the AFM. All maneuvers should be accomplished smoothly, accurately, and without sustained nuisance oscillation.
(f) In addition to performance of in tended function of each mode examined, the automatic pilot system should cause no sustained nuisance oscillations, undue control activity, or sudden large attitude changes, especially when configuration or power changes are taking place.
(g) When automatic functions are provided that may be used with these systems
(e.g., automatic throttle control or yaw damper, etc.), and use of the system is permitted with any of these functions inoperative, the system should comply with these flight test requirements with these additional functions operative and inoperative.
(h) In addition to the quick release controls of § 25.1329(d), in order to show compliance with § 25.1309, an alternative means of disengagement for the automatic pilot should be provided. It should be readily accessible in flight.
(i) It should be possible to disengage the automatic pilot system at any time without unacceptable out-of-trim forces.
(j) The controls, indicators, and warnings should be designed to minimize crew errors. Mode and system malfunction indications should be presented in a manner compatible with the procedures and assigned-tasks of the flightcrew. The indications should be grouped in a logical and consistent manner, and be visible under all lighting conditions.
(k) The means provided to comply with § 25.1329(h) should also give an appropriate indication when there is an inadvertent change or disengagement of a mode.
(2) Takeoff Mode .
(a) AC 25-15, “Approval of Flight Management Systems in Transport Category Airplanes,” contains some basic criteria and airspeed tolerance criteria that should be included in the flight evaluation of systems containing takeoff guidance. Additional testing considerations, which should be considered for inclusion in the flight test plan, are presented in paragraphs (b) through (s), below.
(b) The roll axis control law will probably be a heading hold or roll attitudehold/wings-level mode. The performance of the roll axis control function is considered to be obvious during the normal course of flight testing and, therefore, is not addressed with specific test requirements herein. The pitch axis control law discussed below represents a system providing computed command guidance (flight director, HGS, and automatic pilot after liftoff) representing the reference speed to be flown during second segment climb-out.
(c) For a normal all-engines-operating (AEO) takeoff, the reference speed should be the normal AEO initial takeoff climb speed selected by the applicant (typically V2+10 knots).
In the event that an engine failure occurs during the takeoff maneuver, at an airspeed greater than V1, the reference speed should be:
1 V 2 if an engine failure occurs prior to V2.
2 Existing speed if engine failure occurs between V 2 and the AEO initial climb speed.
3 The AEO initial takeoff climb speed for all other conditions, if appropriate.
(d) Evaluate takeoff performance with different flap settings and different loading conditions (gross weight and center of gravity) to cover the various flight conditions. Use normal rotation and cleanup procedures unless noted otherwise.
(e) Evaluate abnormal/simulated failure conditions including:
1 Late rotation.
2 Early rotation.
3 Simulated engine failure at V 1.
4 Simulated engine failure at V 2+20 knots.
5 Simulated engine failure at V 2+5 knots.
(f) Evaluate the ability of the autopilot to command the correct speed for the conditions in the testing described in paragraphs (c), (d), and (e) above.
(g) If a variable takeoff thrust derate feature is provided, perform takeoff demonstrations using up to the maximum derated thrust level for which approval is to be granted.
(h) If an ATTCS is included on the ai rplane, evaluate its performance of intended function in conjunction with the above.
(i) Conduct takeoffs with and without, as applicable, automatic throttles, yaw damper, automatic brake system, control wheel steering, and any other automated equipment configurations, and combinations thereof, for which the applicant is seeking certification approval. Include all sensor and instrument switching combinations.
(j) Evaluate Slow/Fast display behavior, if applicable, in conjunction with these takeoff tests.
(k) Evaluate the autopilot behavior with both series and parallel rudder (if available) configurations, particularly in conjunction with simulated engine failure tests.
(l) Evaluate “fade-out” of parallel rudder (with simulated engine failure) during autopilot takeoff mode transition to automatic altitude captures (if this feature is available).
(m) Evaluate pitch axis takeoff behavior during simultaneous heading select turns during climb-out.
(n) Evaluate mode transition from ta keoff to other approved modes using flight director, HGS, and automatic pilot guidance (i.e., from pitch axis takeoff mode to IAS Hold or Vertical Speed or Altitude Hold or vertical navigation with autopilot operation (VNAV); from roll axis takeoff mode to Heading Select or VOR or lateral navigation with autopilot operation (LNAV), etc.).
(o) Evaluate all reversion modes appropriate to the design. Most speed command takeoff modes will be fail-passive (or “fail-soft”) and will probably contain reversion modes.
Consult the system description, failure modes and effects analysis (FMEA), and the responsible systems engineer for definition of the appropriate evaluation. This may be accomplished in an appropriate simulator environment.
(p) Evaluate adequacy of mode annunciations and controls and displays associated with takeoff mode operation.
(q) In addition to evaluating speed performance to the criteria shown in Table 2 of AC 25-15, “Approval of Flight Management Systems in Transport Category Airplanes,” evaluate command bar flyability, dynamics of the airplane while following the commanded guidance, initial pitch target presented during and immediately following rotation, etc.
(r) Evaluate takeoff pitch attitude limit, if applicable, for compatibility and command harmony.
(s) Quantitative time history data should be derived with an on-board data system to facilitate post-flight evaluation of performance and to provide type inspection report (TIR) records.
(3) Climb, Cruise, Descent, and Holding Modes .
(a) The acceptability of the autopilot system performance may be based on subjective judgment, taking into account the experience acquired from similar equipment and the general behavior of the airplane. The acceptable performance may vary according to airplane type and model.
(b) Examination of the following modes are considered appropriate for inclusion in this section:
Pitch Attitude Hold Roll Attitude Hold Vertical Speed Hold/Select Heading Hold/Select Altitude Hold/Select VOR Navigation IAS Hold/Select Lateral Navigation Mach Hold/Select Backcourse Level Change Non-Precision Approach Vertical Navigation VOR Turbulence (Penetration) Localizer (only) Area Navigation
Evaluation/approval of these modes generally will not require derivation of quantitative flight data.
(c) Special characteristics associated with some specific modes worthy of note are:
1 Operation of the system should not result in performance for which the pilot would be cited during a check ride (i.e., exceeding a speed target of 250 knots by more than 5 knots, if appropriate, during operations below 10,000 feet altitude, or overshooting a target altitude by more than 100 feet during capture of the pre-selected altitude).
2 Where the automatic pilot system has the ability to acquire and maintain a pre-selected altitude, it should be shown in particular that:
(aa) With the autopilot engaged, if the pilot fails to advance the throttles following an altitude acquisition from the descent, ensure that the airplane exhibits no hazardous characteristics, if recovery action is taken within a reasonable period after the onset of stall warning, or other reliable warning. The autopilot should disconnect at stall warning unless proven safe beyond this point.
NOTE: Compliance with this provision need not be demonstrated if adequate means are provided to prevent such an error.
(bb) Resetting the datum pressure or the selected altitude at any time during altitude capture should not result in hazardous maneuvers.
3 The altitude hold mode should be evaluated:
(aa) During turning flight.
(bb) During accelerating and decelerating flight.
(cc) With engagement of the mode from climbing and descending flight conditions.
4 Thoroughly explore transition of VOR navigation from cruise/holding to the approach mode. Include appropriate procedures in the AFM if special procedures or limitations are deemed to be necessary.
(4) Go-Around Mode .
(a) AC 25-15 also contains some basic criteria and airspeed tolerance criteria for go-around systems that should be included in the flight evaluation. Also, AC 120-28D, “Criteria for Approval of Category III Weather Minima for Takeoff, Landing, and Rollout,” contains additional go-around criteria pertinent to this flight evaluation, if the airplane is to be approved for CAT III operation.
(b) As with the takeoff mode, the go-around roll axis control law will probably be a heading-hold or roll attitude-hold/wings-level mode and will not be addressed herein. The