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
253 pitch axis control law may be an angle-of-attack control law system, but most modern systems will have control laws for speed control. The speed should be compatible with that used for a manually controlled go-around without guidance.
(c) Under no circumstances should the system command a go-around speed less than 1.13 VSR1 for the go-around flap setting (which equates to approximately 1.3 VSR0 + 3 knots for the landing flap setting) or other firmly established minimum control speed (e.g., VMCL) established for that particular airplane. In making this evaluation, consideration should be given to the autopilot roll axis control law and its effect on minimum control speeds; if the autopilot employs a wings-level control law, the minimum control speeds will be considerably increased over the pilot flying values that are determined with some degree of bank into the operating engine(s).
(d) Evaluate go-around performance with normal conditions for each landing flap setting through a full range of gross weight and center of gravity conditions. Use normal cleanup procedures to the appropriate approach climb flap setting.
(e) Evaluate go-around performance with simulated engine failure conditions occurring:
1 Just prior to go-around initiation;
2 Just after go-around initiation; and
3 During climb-out, both before and after initiation of altitude capture mode.
(f) If engine-out approach/landing approval is sought by the applicant, then evaluate go-around performance initiated from an engine-out approach.
(g) It should be demonstrated that the autopilot will command the correct speed to accomplish a go-around for the conditions in the testing described in paragraphs (d), (e), and
(f), above.
(h) Flight path control following an engine failure during go-around should not require exceptional piloting skill or alertness.
(i) Conduct go-arounds with and without, as applicable, automatic throttles, yaw damper, automatic ground spoilers, control wheel steering, and any other equipment configurations, and combinations thereof, for which the applicant is seeking certification credit.
Include all sensor and instrument switching combinations.
(j) Evaluate Slow/Fast display behavior, if applicable, in conjunction with these go-around 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 go-around mode transition to automatic altitude captures (if this feature is available).
(m) Verify that when automatic go-around is engaged, subsequent momentary ground contact will not cause its disengagement or a mode reversion out of go-around.
(n) Structure the test program such that altitude loss during the go-around maneuver can be determined as a function of height above the ground at go-around initiation.
Altitude loss during the maneuver is directly proportional to the rate of descent on initiation of go-around.
(o) Go-around altitude loss is defined as the difference in altitude between the selection of the go-around mode and the minimum height above the ground achieved during the maneuver.
(p) If the airplane is to be approved for Category III operations, the go-around altitude loss information should be included in the AFM (see appendix 6 of AC 120-28D). The preferred format for presentation of this information in the AFM is as a cross-plot of altitude loss (ordinate of the crossplot) versus altitude of mode initiation (abscissa of the crossplot) as shown below.
Figure 181-1 Altitude Loss versus Altitude Go-Around Mode
EXAMPLE ALTITUDE LOSS DURING GO-AROUND (FT) go-around initiated at 50 feet is The expected altitude loss during between 38 feet and 45 feet.
50 40 These data are derived from tests to 30 determine altitude losses during go arounds.
They were conducted with the autopilot 20 engaged, with all engines functional and with simulated engine failures, and at all 10 approved flap settings. There is essentially no difference in results, regardless of aircraft configuration.
20 40 60 80 100
ALTITUDE AT GO-AROUND INITIATION (FT)
(q) Evaluate the go-around maneuver under conditions with automatic spoilers armed and go-around initiated at a low altitude, which results in wheel contact with the runway surface. Verify that momentary deployment of the ground spoilers under these conditions is acceptable.
(r) In addition to evaluation of 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, the dynamics of the airplane while following the commanded guidance, etc. The control action and flight path during the initial rotation should not be significantly different from those of a manually controlled go-around without command guidance.
(s) The go-around from any point on the approach to touchdown should not require exceptional piloting skill, alertness, or strength, and should ensure that the airplane remains above the obstacle limitation surface specified in AC 120-29A, "Criteria for Approval of Category I and Category II Weather Minima for Approach."
(t) Evaluate pitch axis go-around mode behavior with simultaneous heading select turns during climb-out.