(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 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) In accordance with § 25.1310, for probable operating combinations of required electrical installations, verify that the following power loads are provided for probable durations:
(a) Loads connected to the system with the system functioning normally;
(b) Essential loads after failure of any one prime mover, power, converter, or energy storage device;
(c) Essential loads after failure of one engine on a two-engine airplane;
(d) Essential loads after failure of two engines on airplanes with three or more engines;
(e) Essential loads for which an alternate source of power is required, after any failure or malfunction in any one power supply system, distribution system, or other utilization system.
(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.
(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/thrust. 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) If an alternate system is provided, it may be calibrated against either the reference system or the airplane’s system.
(d) An acceptable means of compliance when demonstrating a perceptible speed change between 1.23 VSR to stall warning speed (§ 23.1223(d)) is for the rate of change of IAS with CAS to be not less than 0.75.
(e) 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.
(f) 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.
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
255 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) Power or Thrust - as required.
b. Procedures.
(1) Any one or any desired combination of the procedures in subparagraphs (2) or (3) 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 1.23 VSR. The highest speed should not exceed VMO/MMO. 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) 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.
(3) 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.
(4) 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.
(5) 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
257 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, or vice versa, if both use the same static vent, provided any error in total pressure over the angle-of-attack range is taken into account.
b. Procedures. None.
179. Pitot Heat Indication Systems - § 25.1326. [Reserved]
180. Magnetic Direction Indicator - § 25.1327. [Reserved]
181. Flight Guidance System - § 25.1329.
a. Explanation.
(1) On most modern airplane installations, the autopilot and flight director are integrated into a single computer, use common control laws, and are identified as a flight guidance system (FGS). The FGS is primarily intended to assist the flightcrew in the basic control and tactical guidance of the airplane. The system may also provide workload relief to the pilots and provide a means to fly a flight path more accurately to support specific operational requirements, such as reduced vertical separation minimum (RVSM) or required navigation performance (RNP). To perform these functions, the FGS typically includes an integrated autopilot, autothrust, and flight director. The FGS’s functions also include flight deck alerting, status, mode annunciations (instrument displays), and associated information displayed to the flightcrew for situation awareness. Also included are those functions necessary to provide guidance and control with an approach and landing system, such as an instrument landing system (ILS), microwave landing system (MLS), global navigation satellite system (GNSS), or a GNSS landing system (GLS). 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. (NOTE:
For applicants considering type certification of a HUD system, we recommend that the flight technical error be characterized during certification testing. This data will be required if the applicant chooses to seek operational credit for advanced performance-based navigation applications.
(2) For the purposes of this AC, the FGS includes all the equipment necessary to accomplish the FGS function, including the sensors, computers, power supplies, servomotors/actuators, and associated wiring. It also includes any indications and controllers necessary for the pilot to manage and supervise the system.
(3) A flight test program should be established that confirms the performance of the FGS for the modes of operation and the operational capabilities supported by its design. The operational implications of certain failures and failure conditions may require flight evaluation.
Also, the pilot interface with FGS controls and displays in the cockpit will need to be assessed.
Some aspects of the FGS design may be validated by laboratory test and/or simulator evaluation.
(4) The scope of the flight demonstration program will depend on the operational capability being provided, including any new and novel features. Early coordination with the FAA is recommended to reduce certification risks associated with the flight demonstration program. The intent of the flight demonstration program is to confirm that the operation of the FGS is consistent with its use for the intended flight operations of the airplane type and configuration. The modes of the FGS should be demonstrated in representative airplane configurations and under a representative range of flight conditions.
(5) Additional guidance material related to FGS operation, testing, and approval is contained in the following FAA ACs:
(a) AC 25-15, “Approval of Flight Management Systems in Transport Category Airplanes,” dated November 20, 1989.
(b) AC 25.1329-1B Change 1,“Approval of Flight Guidance Systems,” dated October 16, 2012.
(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 the intended function(s) and safe operation of the system. Flight test evaluations supplement analysis, laboratory testing, and simulator testing for showing compliance with the applicable part 25 requirements.
(b) The requirement for demonstrating safe operation should include evaluation of FGS failure conditions that a system safety assessment identifies as needing to be validated by flight test.
(c) The FGS should be installed and adjusted so that the system tolerances established during certification tests can be maintained in normal operation. This may be addressed by conducting flight tests at the extremes of the tolerances. Those tests conducted to determine that the FGS will adequately control the airplane should establish the lower limit, and those tests to determine that the FGS 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) 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 is included in the AFM. All maneuvers should be accomplished smoothly, accurately, and without sustained nuisance oscillation.
(e) In addition to performance of intended function of each mode examined, the FGS should cause no sustained nuisance oscillations, undue control activity, or sudden large attitude changes, especially when configuration or power/thrust changes are taking place.
(f) When use of the FGS is permitted with any of its functions inoperative (e.g., autothrust, yaw damper, etc.), the system should be evaluated with these functions both operative and inoperative.
(2) FGS Protection Features.
(a) Section 25.1329(h) requires that when the flight guidance system is in use, a means must be provided to avoid excursions beyond an acceptable margin from the speed range of the normal flight envelope. The FGS itself may contain protection features to aid the flight crew in assuring that the boundaries of the flight envelope or operational limits are not exceeded.
The means to alert the flightcrew to a boundary or for the system to intervene may vary, but certain operational scenarios can be used to assess the performance of the system in providing the protection function. The procedures in the following paragraphs can be used to evaluate the protection features as required by § 25.1329(h), regardless of whether they are included in the FGS itself, or invoked by other means.
(b) Low Speed Protection. Low speed protection is intended to prevent loss of speed leading to an unsafe condition. If the FGS remains in the existing mode, a suitable alert should be provided to annunciate the low speed condition. In this case, note the pilot response to the alert and the recovery actions taken to maintain the desired vertical path and to accelerate back to the desired speed. The following scenarios should be considered when evaluating the low speed protection means that must be present when the FGS is in use:
1 High Altitude Cruise Evaluation.
(aa) At high altitude at normal cruise speed, engage the FGS into an Altitude Hold mode and a heading or lateral navigation (LNAV) mode.
260 (bb) Engage the autothrust into a speed mode.
(cc) Manually reduce one engine to idle power or thrust.
(dd) As the airspeed decreases, observe the FGS behavior in maintaining altitude and heading/course.
(ee) When the low speed protection feature becomes active, note the airspeed and the associated aural and visual alerts including possible mode change annunciations for acceptable operation.
2 Altitude Capture Evaluation at Low Altitude.
(aa) At a reasonably low altitude (e.g., approximately 3000 feet above MSL where terrain permits) and at 250 knots, engage the FGS into Altitude Hold and a heading or LNAV mode.
(bb) Engage the autothrust into a speed mode.
(cc) Set the altitude pre-selector to 5000 feet above the current altitude.
(dd) Make a flight level change to the selected altitude feet with a 250 knots climb at maximum climb power or thrust.
(ee) When the FGS first enters the Altitude Capture mode, reduce thrust/power on one engine to idle.
(ff) As the airspeed decreases, observe the airplane trajectory and behavior.
(gg) When the low speed protection condition becomes active, note the airspeed and the associated aural and visual alerts including possible mode change annunciations for acceptable operation.
3 High Vertical Speed Evaluation.
(aa) Engage the FGS in the Vertical Speed mode with a very high rate of climb.
(bb) Set the thrust/power to a value that will cause the airplane to decelerate at approximately 1 knot per second.
(cc) As the airspeed decreases, observe the airplane trajectory and behavior.
(dd) When the low speed protection condition becomes active, note the
261 airspeed and the associated aural and visual alerts including possible mode change annunciations for acceptable operation.
4 Approach Evaluation.
(aa) Conduct an instrument approach with vertical path reference.
(bb) Couple the FGS to the localizer and glideslope (or LNAV/VNAV, etc.).
(cc) Cross the final approach fix/outer marker at a reasonably high speed at idle thrust/power until low speed protection activates.
(dd) As the airspeed decreases, observe the airplane trajectory and behavior.
(ee) When the low speed protection becomes active, note the airspeed and the associated aural and visual alerts including possible mode change annunciation for acceptable operation.
(ff) Note the pilot response to the alert and the recovery actions taken to recover to the desired vertical path and the re-capture to that path and the acceleration back to the desired approach speed.
(c) High Speed Protection. High speed protection, which may either be included in the FGS or provided by other means, is intended to prevent a gain in airspeed leading to an unsafe condition. If the FGS remains in the existing mode with reversion to high speed protection, a suitable alert should be provided to annunciate the high speed condition. In this case, note the pilot response to the alert and the recovery actions taken to maintain the desired vertical path and to decelerate back to the desired speed. The following scenarios should be considered when evaluating the means used to provide protection from high speed excursions while the FGS is in use, as required by § 25.1329(h):
1 High Altitude Level Flight Evaluation with Autothrust Function.
(aa) Select autothrust off, if an automatic wake-up function is provided;
otherwise, select autothrust on.
(bb) Engage the FGS in Altitude Hold.
(cc) Select a power or thrust level that will result in an acceleration that would cause, without intervention (either automatic or manual), a speed/Mach beyond VMO/MMO.
(dd) As the airspeed increases, observe the behavior of the high-speed protection condition and any autothrust reactivation and thrust/power reduction, as applicable.
(ee) Assess the performance of the FGS to control the airspeed to VMO/MMO, or other appropriate speed.
2 High Altitude Level Flight Evaluation Without Autothrust Function.
(aa) Set a power/thrust level that will result in acceleration that would cause, without intervention (either automatic or manual), a speed/Mach beyond VMO/MMO.
(bb) As the airspeed increases, observe the basic airplane overspeed warning activate.
(cc) Observe activation and effectiveness of the high speed protection means and note any FGS indications and behavior.
(dd) Maintain the existing power or thrust level and observe the airplane depart the selected altitude.
(ee) After sufficient time has elapsed to verify and record FGS behavior, reduce the thrust/power as necessary to cause the airplane to begin a descent.
(ff) Observe the FGS behavior during the descent and subsequent altitude capture at the original selected altitude.
3 High Altitude Descending Flight Evaluation With Autothrust Function.
(aa) Select autothrust off (with automatic wake-up function) with power or thrust set to maintain airspeed 10 percent below VMO/MMO with the FGS engaged in Altitude Hold.
(bb) Select Vertical Speed mode with a vertical speed that will result in an acceleration that would cause, without intervention (either automatic or manual), a speed/Mach beyond VMO/MMO.
(cc) As the airspeed increases observe the autothrust function reactivate and reduce power or thrust towards idle.
(dd) Observe the activation and effectiveness of the high speed protection means and note any FGS indications and behavior.
(3) FGS 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,
263 which should be considered for inclusion in the flight test plan, are presented in paragraphs (b) through (s), below.
(b) There are no test conditions specifically identified for evaluating the roll axis control function. The performance of this function should be evaluated during the normal course of flight testing. The test conditions for the pitch axis control function identified below is for a system providing computed command guidance (flight director, Head Up Guidance System (HGS), and automatic pilot after liftoff) to achieve and maintain a 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 VEF, the reference speed should be:
1 V2 if an engine failure occurs prior to V2.
2 The existing speed if engine failure occurs between V2 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 c.g.) 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 VEF.
4 Simulated engine failure at V2+20 knots.
5 Simulated engine failure at V2+5 knots.
(f) Evaluate the ability of the FGS to command the correct speed for the conditions in the testing described in paragraphs (c), (d), and (e) above.
(g) If a variable takeoff power or thrust derate or reduced thrust (power) feature is provided, perform takeoff demonstrations to cover the range of derated/reduced power or thrust levels for which approval is requested.