3. With engagement of the mode from climbing and descending flight conditions.
33.7.5.4.4 Thoroughly explore transition of VOR navigation from cruise/holding to the approach mode. Include appropriate procedures or limitations in the AFM if special procedures or limitations are deemed to be necessary.
33.7.6 Procedures: FGS Go-Around Mode.
33.7.6.1 AC 25-15 contains some basic criteria and airspeed tolerance criteria for go-around systems that should be included in the flight evaluation. Also, AC 120-28D contains additional go-around criteria pertinent to this flight evaluation if the airplane is to be approved for CAT III operation.
33.7.6.2 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. The 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.
33.7.6.3 The system should not command a go-around speed less than 1.13 VSR1 for the go-around flap setting or the 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 may be considerably higher than the AFM values that are determined with some degree of bank into the operating engine(s).
33.7.6.4 Evaluate go-around performance with normal conditions for each landing flap setting through a full range of gross weight and CG conditions. Use normal cleanup procedures to the appropriate go-around flap setting.
33.7.6.5 Evaluate go-around performance with simulated engine failure conditions occurring:
Just prior to go-around initiation, Just after go-around initiation, and During climb-out, both before and after initiation of altitude capture mode.
33.7.6.6 If engine-out approach/landing approval is sought by the applicant, then evaluate go-around performance initiated from an engine-out approach.
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33.7.6.7 It should be demonstrated that the FGS will command the correct speed to accomplish a go-around for the conditions in the testing described in paragraphs 33.7.6.4, 33.7.6.5, and 33.7.6.6 above.
33.7.6.8 Flight path control following an engine failure during go-around should not require exceptional piloting skill or alertness.
33.7.6.9 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.
33.7.6.10 Evaluate Slow/Fast display behavior, if applicable, in conjunction with these go-around tests.
33.7.6.11 Evaluate the autopilot behavior with both series and parallel rudder (if available) configurations, particularly in conjunction with simulated engine failure tests.
33.7.6.12 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).
33.7.6.13 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.
33.7.6.14 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 when the go-around is initiated.
33.7.6.15 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.
33.7.6.16 Go-around altitude loss information should be included in the AFM, especially if the airplane is to be approved for low visibility (Category I, II, or III) approaches. An example of how such information has been presented in the AFM is shown in figure 33-1 below.
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Figure 33-1. Altitude Loss versus Altitude at Go-Around Mode Initiation
33.7.6.17 Evaluate the go-around maneuver under conditions with automatic spoilers armed and go-around initiated at an altitude low enough to result in wheel contact with the runway surface. Verify that momentary deployment of the ground spoilers under these conditions is acceptable.
33.7.6.18 In addition to evaluating speed performance to the criteria shown in Table 2 of AC 25-15 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.
33.7.6.19 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.
33.7.6.20 Evaluate pitch axis go-around mode behavior with simultaneous heading select turns during climb-out.
33.7.6.21 Evaluate mode transitions from go-around to other approved modes using flight director, HGS, and automatic pilot guidance (i.e., from pitch axis go-around mode to IAS Hold, Vertical Speed, Altitude Hold, or VNAV;
from roll axis go-around mode to Heading Select VOR, or LNAV, etc.).
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33.7.6.22 Evaluate all reversion modes appropriate to the design. Most speed command go-around modes will be fail-passive (or fail-soft) and will probably contain reversion modes. Consult the system description, the FMEA, and the responsible system engineer for definition of the appropriate evaluation. It may be possible to perform this testing in a simulator.
33.7.6.23 Evaluate adequacy of mode annunciations and controls and displays associated with the go-around maneuver.
33.7.6.24 Evaluate the go-around pitch attitude limit (if applicable) for compatibility and command harmony.
33.7.6.25 Fixed Pitch Go-Around.
Missed approaches using the calibrated attitude markings on the Attitude Director Indicator (ADI) should be demonstrated to be an acceptable alternative technique for low visibility operations (CAT II & CAT III), in the event that automatic or flight director go-around modes are not available. A sufficient sample size should be collected to establish the transitional altitude loss. The target attitude for go-around (13° to 17° is usually the target) should be established by the applicant. Use normal cleanup procedures. Simulate various failure conditions that would force the pilot to revert to this configuration. (e.g., simultaneous loss of autoland and go-around computation).
33.7.6.26 Quantitative time-history data should be derived with an on-board data system to facilitate post-flight evaluation of performance and to provide TIR records.
33.7.7 Procedures: FGS Instrument Landing System (ILS) Approach Mode.
33.7.7.1 A complete and comprehensive description of flight test evaluation for automatic pilot and flight director system ILS approach mode approval, including HUDs, is contained in AC 120-29A. Further guidance is provided below:
33.7.7.2 For airworthiness approval to Category I (CAT I) minimums:
33.7.7.2.1 Conduct a series of approaches (usually 4 or more) on Type I rated ILS beams to a radio altitude of 160 feet. (Twenty percent below the CAT I decision height of 200 feet).
33.7.7.2.2 Conduct the approaches with and without automatic throttles, with and without yaw damper, alternate flap setting and so forth, for all configurations and combinations of equipment for which the applicant seeks approval.
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33.7.7.2.3 At least three Type I beams should be included in the evaluation, one of which should exhibit very noisy localizer and glideslope characteristics.
33.7.7.2.4 Failure modes/conditions described in paragraph 33.7.8.6 of this AC, appropriate to ILS approach modes, should be conducted.
33.7.7.2.5 The definition of a successful approach is one that positions the airplane at the decision height (DH) such that the airplane can be safely landed without exceptional piloting skill or strength.
33.7.7.3 For airworthiness approval to Category II (CAT II) minimums for each candidate system (automatic pilot, flight director, and HGS):
33.7.7.3.1 Conduct a series of approaches on Type II rated ILS beams to a radio altitude of 100 feet.
33.7.7.3.2 For initial system approval, approximately 20 approaches (total for each affected system) are required to examine each configuration/combinations of equipment for which the applicant seeks credit.
33.7.7.3.3 At least three Type II beams should be included in the evaluation.
33.7.7.3.4 When relatively minor changes are made to the approach control laws or when a different display (i.e., flight director instrument) is used with the system, approximately 9 approaches (total for each affected system) have been found to be sufficient (i.e., 3 each on 3 different ILS beams).
33.7.7.3.5 The approaches should be made in conditions chosen to show that the performance is satisfactory within the permitted extremes of weight, CG, wind speed, localizer capture angles, glideslope captures from below and above (if appropriate) the beam, captures at various ranges from runway threshold, etc.
33.7.7.4 If approval is sought for ILS approaches initiated with one engine inoperative, and with the airplane trimmed at the point of glide path intercept, the automatic flight control system should be capable of conducting the approach without further manual trimming. For airplanes with three or more engines, the loss of a second critical engine should not cause a rate of lateral deviation from the ILS course greater than 3° per second (averaged over a 5-second period) or produce hazardous attitudes.
33.7.7.5 Unless it is shown that failure of the automatic pilot system to disengage during the approach, when the pilot operates the quick release control on the control wheel, is improbable, then it should be demonstrated that the pilot can control the airplane manually without operating any of the other disengagement controls.
33.7.7.6 It is possible that many of the above tests can be performed in a simulator.
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33.7.8 Procedures: FGS Control Wheel Steering (CWS).
33.7.8.1 It should be possible for the pilot to overpower the automatic pilot system, and achieve the maximum available control surface deflection, without using forces that exceed the pilot control force limits specified in § 25.143(d).
33.7.8.2 The maximum bank and pitch attitudes that could be achieved without overpowering the automatic pilot system should be limited to those necessary for the normal operation of the airplane. Typically, these attitudes are ±35° in roll and +20° to -10° in pitch.
33.7.8.3 It should be possible to carry out all normal maneuvers and to counter normal changes of trim due to change of configuration etc., within the range of flight conditions in which control wheel steering may be used, without encountering excessive discontinuities in control force that might adversely affect the flight path.
33.7.8.4 The stall and stall recovery characteristics of the airplane should remain acceptable with control wheel steering in use.
33.7.8.5 In showing compliance with § 25.143(f), account should be taken of such adjustment to trim as may be carried out by the automatic pilot system in the course of maneuvers that can reasonably be expected. Some alleviation may be acceptable in the case of unusually prolonged maneuvers, provided the reduced control forces would not be hazardous.
33.7.8.6 If the use of this mode for takeoff and landing is to be permitted, it should be shown that:
Sufficient control, both in amplitude and rate, is available without encountering force discontinuities;
Reasonable mishandling is not hazardous (e.g., engaging the autopilot while the pitch or roll controls are held in an out-of-trim position); and Runaway rates and control forces are such that the pilot can readily overpower the automatic pilot with no significant error in flight path.
33.7.9 Procedures: Environmental Conditions.
33.7.9.1 Some environmental conditions have created operational problems during FGS operations. The flight demonstration program should expose the FGS to a range of environmental conditions as the opportunity presents itself.
These include winds, mountain-wave, turbulence, icing, etc. However, some specific test conditions may have to be flown to find operational conditions that are not readily achieved during the normal flight test program 33-19
AC 25-7D CHG 1
33.7.9.2 FGS use in icing conditions may mask the onset of an airplane condition
(e.g., out of trim condition) that would present the pilot with handling difficulties if the FGS is disengaged, particularly if the FGS suddenly disengages automatically. The opportunity should be taken during the flight test program to evaluate the FGS during natural icing conditions, including during shedding of the ice, as applicable. The operation of the FGS should also be evaluated during basic airplane performance and handling qualities compliance flight tests with simulated ice accretions installed. (See AC 25-25A for examples of a flight test program for evaluating airplane performance and handling qualities in icing conditions and for guidance on ice accretions to use for flight testing.) The following test conditions should also be considered for evaluating FGS performance in icing conditions:
33.7.9.2.1 Low Speed Protection.
1. With all high lift devices retracted, slow down at a rate not to exceed 1 knot per second until either the autopilot automatically disengages (with an associated alert) or a low speed protection function engages (which may be a low speed alert).
2. Recovery should be initiated no less than three seconds after the onset of an appropriate alert.
3. The airplane should exhibit no hazardous characteristics.
33.7.9.2.2 Coupled Approach.
If the autopilot has the ability to fly a coupled instrument approach and go-around, the following test conditions should be evaluated:
1. Instrument approach, using all normal flap selections.
2. Go-around, using all normal flap selections.
3. Glideslope capture from above the glidepath.
33.7.9.2.3 If the airplane accretes or sheds ice asymmetrically, it should be possible to disengage the autopilot at any time without unacceptable out-of-trim forces.
33.7.9.2.4 General maneuverability should be assessed, including normal turns, and the maximum angle of bank commanded by the FGS in one direction followed by a rapid reversal of command reference to the maximum FGS angle of bank in the other direction.
33.7.10 Procedures: Failure Modes/Malfunction Tests.
33.7.10.1 General.
33.7.10.1.1 Failure conditions of the system should be simulated in such a manner as to represent the overall effect, and worse case effect, of each failure 33-20
condition about all axes. The test method for most failure conditions will require some type of fault simulation technique with controls that provide for controlled insertion and removal of the type of fault. The insertion point will typically be at a major control or guidance point on the airplane (for example, a control surface command, guidance command, or power/thrust command).
33.7.10.1.2 Investigations should include the effects of any failure conditions identified for validation by a system safety assessment conducted to show compliance with § 25.1309(d).
33.7.10.1.3 The safety assessment process may identify vulnerabilities to failure conditions involving hardovers, slowovers, or oscillatory behavior. The various types of effect will result in differing airplane responses and cues to alert the flightcrew to the failure condition. The recognition point should be that at which a line pilot in non-visual conditions may be expected to recognize the need to act. Recognition of the malfunction may be through the behavior of the airplane or a reliable alerting system. The recognition point should be identified by the test pilot. Control column or wheel movements alone should not be used for recognition.
33.7.10.1.4 After the recognition point, action by the test pilot should be delayed to simulate the time it would take for a line pilot to take control after recognizing the need for action. The test condition is considered completed when a stable state is reached as determined by the test pilot. A three-second delay added to the measured time increment between pilot recognition of an automatic pilot malfunction and pilot corrective action has been considered acceptable for climb, cruise, and descent phases of flight. A one-second delay is considered appropriate for flight phases where the crew is expected to be closely monitoring FGS control inputs, such as during approach, and for FGS use during takeoff from shortly after liftoff through flap retraction. For control wheel steering (CWS) mode of operation during takeoff, go-around, and landing, and for autoland and goaround mode hardovers, no delay time need be applied.
33.7.10.1.5 Satisfactory airplane response to autopilot hardovers should be shown throughout the entire certificated airspeed/altitude flight envelope. Since loading limitations of the test airplane preclude investigation of the entire flight envelope for which certification is requested, simulation results may be used to validate response in the unreachable CG versus weight combinations.
33.7.10.1.6 The nominal worst-case weight/CG combinations achievable by the test airplane, as predicted by simulation results, should be flight tested. The justification for selection of the weight/CG combinations proposed for 33-21
testing should be submitted for FAA review prior to type inspection authorization (TIA) issuance.
33.7.10.1.7 The analysis should also present a comprehensive exploration of airplane “g” responses throughout the weight-CG-airspeed-altitude flight envelope, including worst case and envelope boundary conditions. The analysis should include flight test and simulator responses adjusted for the presence of worst-case system tolerance effects.
33.7.10.1.8 During recovery from an automatic pilot system malfunction, the pilot may overpower the autopilot or disengage it. The pilot should be able to return the airplane to its normal flight attitude under full manual control, without exceeding the loads or speed limits appropriate to the flight condition, without engaging in any dangerous maneuver during recovery, and without control forces exceeding the values given in § 25.143(d).
33.7.10.1.9 If an autothrottle is installed, the malfunctions should be examined with and without the autothrottle operating.
33.7.10.1.10 The airplane should be instrumented such that parameters appropriate to the tests are recorded (e.g., normal acceleration, airspeed, altitude, pitch and roll attitude, automatic pilot engagement discrete).
33.7.10.2 Oscillatory Tests.
33.7.10.2.1 An investigation should be made to determine the effects of an oscillatory signal of sufficient amplitude to saturate the servo amplifier of each device that can move a control surface. The investigation should cover the range of frequencies that can be induced by a malfunction of the automatic pilot system and systems functionally connected to it, including an open circuit in a feedback loop. The investigated frequency range should include the highest frequency that results in apparent movement of the system driving the control surface to the lowest elastic or rigid body response frequency of the airplane. Frequencies less than 0.2 Hz may, however, be excluded from consideration. The investigation should also cover the normal speed and configuration ranges of the airplane. The results of this investigation should show that the peak loads imposed on the parts of the airplane by the application of the oscillatory signal are within the limit loads for these parts. Flight guidance systems that contain integral performance envelope limiting cut-out functions may be exempt from this requirement, provided the monitor is demonstrated to have sufficient integrity.
33.7.10.2.2 The investigation may be accomplished largely through analysis with sufficient flight data to verify the analytical studies, or largely through flight tests with analytical studies extending the flight data to the conditions that impose the highest percentage of limit load to the parts.
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33.7.10.2.3 When flight tests are conducted in which the signal frequency is continuously swept through a range, the rate of frequency change should be slow enough to permit determination of the amplitude of response of any part under steady frequency oscillation at any critical frequency within the test range.
33.7.10.3 Climb, Cruise and Descent Flight Regimes.
33.7.10.3.1 The more critical of the following should be induced into the FGS. If autothrottles are installed, they should be operating, and vertical gyro mechanical failures should not be considered:
1. A signal about any axis equivalent to the cumulative effect of any single failure, including autotrim, if installed.
2. The combined signals about all affected axes, if multiple axis failures can result from the malfunction of any single component.
33.7.10.3.2 Corrective action should not be initiated until three-seconds after the pilot has become aware, either through the behavior of the airplane or a reliable failure warning system, that a malfunction has occurred.
33.7.10.3.3 The simulated failure and the subsequent corrective action should not result in accelerations normal to the flight path below zero g or above 2 g, speeds beyond VFC/MFC, or result in dangerous dynamic conditions or deviations from the flight path. The positive “g” limitation may be increased up to the positive design limit maneuvering load factor, provided adequate analysis and flight test measurements are conducted to establish that no resultant airplane load is beyond limit loads for the structure, including a critical assessment and consideration of the effects of structural loading parameter variations, (i.e., CG, load distribution, control system variations, etc.). Analysis alone may be used to establish that limit loads are not exceeded where the airplane loads are in the linear range of loading, (i.e., aerodynamic coefficients for the flight condition are adequately established and no significant nonlinear air loading exists). If significant nonlinear effects could exist (e.g., buffet loads), flight loads survey measurements may be necessary to substantiate that the limit loads are not exceeded.
33.7.10.3.4 The power or thrust for climb should be the most critical of that used:
In the performance climb demonstrations;
In the longitudinal stability tests; or For normal operational speeds.
33.7.10.3.5 The altitude loss for the cruise condition is the difference between the observed altitude at the time the malfunction is introduced, and the lowest altitude observed in the recovery maneuver.
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33.7.10.4 Maneuvering Flight.
33.7.10.4.1 Maneuvering flight tests should include turns with the malfunctions introduced when maximum bank angles for normal operation of the system have been established, and in the critical airplane configuration and stages of flight likely to be encountered when using the automatic pilot.
33.7.10.4.2 A one second delay time following pilot recognition of the malfunction, through the behavior of the airplane or a reliable failure warning system, should be used for maneuvering flight malfunction testing.
33.7.10.4.3 The altitude loss, for maneuvering flight testing, is the difference between the observed altitude at the time the malfunction is introduced, and the lowest altitude observed in the recovery maneuver.
33.7.10.5 Approach.
There are two types of approach operations to consider—an approach with vertical path reference and one without vertical path reference. The approach with vertical path reference should be assessed against ground-based criteria using a deviation profile assessment. A height loss assessment should be used for approaches without vertical path reference.
33.7.10.5.1 Fault Demonstration Process.
First determine the worst-case malfunction for each vertical flight path mode that may be used to conduct an approach (e.g., ILS, MLS, GLS, flight management system/RNAV, vertical speed, flight path angle), based on factors such as the following:
Failure conditions identified by the system safety assessment.
System characteristics such as variations in authority or monitor operation.
Mitigation provided by any system alerts.
Aircraft flight characteristics relevant to failure recognition.
33.7.10.5.2 Once the worst-case malfunction has been determined, flight tests of the worst-case malfunction should be flown in representative conditions (for example, coupled to an ILS), with the malfunction being initiated at a safe height. The pilot should not initiate recovery from the malfunction until one second after the recognition point. The delay is intended to simulate the variability in response to effectively a “hands off” condition. It is expected that the pilot will follow through on the controls until the recovery is initiated.
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33.7.10.5.3 Assessment of Approach with Vertical Path Reference.
Figure 33-2 depicts the deviation profile method. The first step is to identify the deviation profile from the worst-case malfunction. The next step is to “slide” the deviation profile down the glidepath, until it is tangential to the 1:29 line or the runway. The failure condition contribution to the minimum use height (MUH)-approach may be determined from the geometry of the airplane wheel height determined by the deviation profile, relative to the 1:29 line intersecting a point 15 feet above the threshold.
Note: The MUH-approach is based on the recovery point for the following reasons: (1) it is assumed that in service the pilot will be “hands off” until the autopilot is disengaged at the MUH in normal operation; (2) the test technique assumes a worst-case based on the pilot being “hands off” from the point of malfunction initiation to the point of recovery; and (3) a failure occurring later in the approach than the point of initiation of the worst-case malfunction described above is therefore assumed to be recovered earlier and in consequence to be less severe.
33.7.10.5.4 Assessment of Approach without Vertical Path Reference.
Figure 33-3 depicts the height loss method. A descent path of 3°, with nominal approach speed, should be used unless the autopilot is to be approved for significantly steeper descents. The vertical height loss is determined by the deviation of the aircraft wheel height relative to the nominal wheel flight path.
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