AC 29-2C, Chg 1 2/12/03 (MANUAL IDENTIFICATION) ATTACHMENT 4
TYPE DESIGN CHANGE
ICA RECOMMENDED PROCEDURES (Continued) Requirement Regulation Location (N/A) The Airworthiness Limitations Section must set forth each mandatory replacement time, structural inspection procedure approved under § 29.571.
A29.4 N/A ( Y ) If the Instructions for Continued Airworthiness consist of multiple documents, the Airworthiness Limitations Section required by this paragraph must include in the principal manual.
A29.4 ATA 4 ( Y ) The Airworthiness Limitations Section must contain a legible statement in a prominent location indicating that the Airworthiness Limitations Section is FAA/AUTHORITY-approved and specifies required maintenance and/or inspections. The exact, required wording of this statement is found in the FAR/JAR.
A29.4 ATA 4 Figure 2 (continued) NOTE: The Airworthiness Limitations Section (ALS) is evaluated and approved by FAA/AUTHORITY.
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(MANUAL IDENTIFICATION)
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AC 29 APPENDIX B. AIRWORTHINESS GUIDANCE FOR ROTORCRAFT INSTRUMENT FLIGHT.
a. Explanation.
(1) Requirements for instrument flight rules (IFR) have been incorporated into Appendix B to part 29, utilizing a regulatory format. Various information from previous interim standards, procedures, test techniques, and acceptable means of compliance for rotorcraft IFR flight is included in the following sections.
(2) Amendment 29-51 made a change to paragraph V. Static Lateral Directional Stability, concurrent with the change to § 29.177, to allow for a small range of sideslip angles (2-3 degrees) for which sideslip angles need not increase steadily with control deflection. The previous rule language stating that directional control position must increase in approximate constant proportion with sideslip angle has been replaced. The intent of this change is that an increase in directional control position must produce an increase in sideslip angle linearly. At greater sideslip angles appropriate to the type rotorcraft, increase in directional control position need not produce a linear increase in sideslip angle but should not become neutral or negative. The change in paragraph VII revised the current requirement to clearly state the requirements to be met in the event of a stability augmentation system (SAS) failure.1
b. Procedures.
(1) General.
(i) The certificated instrument flight envelope may be more restrictive than the visual flight rules (VFR) envelope in terms of weight, CG, speed, altitude, or rate of climb and descent. The approved envelope should be operationally practical such that it does not impose constraints with which the crew has difficulty complying.
(ii) Controllability requirements are to be met from 0.9 VMINI to 1.1 VNE.
Stability requirements must be met where specified. Stability devices are to be designed to allow safe flight following a failure. The evaluating pilot should assure that all equipment and devices installed for IFR, including reasonable failures of that equipment, do not compromise the VFR approval for that rotorcraft. For example, stability system failures can cause loss of swashplate or tail rotor control travel when failed in a hardover condition. If the device remains in the hardover position after the stability system is turned off, control capability may be compromised. Cyclic controllability tests at high speed and at the limiting rearward flight condition, or tail rotor tests in sideward flight at high altitude, may reveal a lower control capability and a more restrictive envelope. In addition, controllability testing should be accomplished with the control rigging set at the most adverse production tolerance for the test condition (e.g., minimum forward swashplate for high speed testing).
1 This paragraph was effective 9/17/2009 as part of the new and changed guidance material for the Performance and Handling Qualities final rule (Amdt. 29-51, 73 FR 11002, Feb. 29, 2008).
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AC 29-2C, Chg 7 2/4/2016
(2) Trim. Compliance with the IFR trim requirement may be met by use of a magnetic brake with a recentering button, an electrically driven trim system activated by a “beeper” type control, or other means, so long as the system does not introduce any objectionable discontinuities in the force gradient or otherwise result in objectionable flight characteristics. Trim release devices should be free of objectionable stick jump.
Electrically driven trim systems should have a smooth change in force with a rate compatible with the normal rotorcraft maneuvers. Only the cyclic trim control must exhibit positive self-centering characteristics. Collective and directional controls are not required to incorporate positive self-centering characteristics. Movement of the trim controls should produce a similar effect on the rotorcraft in a plane parallel to that of the control motion. The control system free play and breakout force must be evaluated to assure a close and direct correlation between control input (force and deflection) and rotorcraft response (pitch, roll, yaw, and heave (vertical motion)), and to permit small, precise changes in flight path. If trim control is provided in a SAS, the control should be of such design and so installed that any failure will not create a hazardous condition. If an inadvertent out-of-trim condition can be developed, its effect on the rotorcraft should be investigated. There should be investigation of these failures or malfunctions as outlined in paragraph b.(7), Stability Augmentation Systems, below. The controls for this trim function should be installed such that, the controls should operate in the plane and with the sense of motion of the rotorcraft. Each control means should have the direction of motion plainly marked thereon or adjacent to the control.
(3) Static Longitudinal Stability.
(i) Positive static longitudinal stability is a key IFR requirement, which assures a self-correcting airspeed response and allows a pilot to recognize any substantial change in speed. The phrase “substantial speed change” as used in Appendix B to part 29, paragraph IV., is normally considered to mean at least a 10 knot departure from trim speed. Such a change in airspeed must be accompanied by a stick force clearly perceptible to the pilot (i.e., a discernible and quantifiable force gradient). Very shallow force gradients can be approved for systems with low deadband and low friction.
Systems with significant friction and deadband require much steeper force gradients to be acceptable. The longitudinal force gradient can be determined by either of two methods. The most commonly used method (applicable only to irreversible control systems) measures the cyclic forces with the rotorcraft on the ground and the rotor stopped (with hydraulic and electric power units if required). The force applied to the cyclic stick and the cyclic stick displacement is measured and a plot of stick force versus displacement in each longitudinal direction is obtained. Following the ground test, the longitudinal static stability tests are conducted in flight as described in section
29.175 of this AC. The cyclic displacement measurements gathered during flight test are then assigned force values from the ground mechanical characteristics test and the force values are cross plotted with the corresponding airspeeds to produce a plot of cyclic force versus airspeed. The trim system should be on during the test and the aircraft trimmed at the trim speed. After each end point, the cyclic should be allowed to slowly return to the trim position. When all the force is released from the cyclic stick and Page Apdx B - 2
the airspeed has stabilized, note the airspeed. An alternate method of determining the longitudinal stick force stability is to measure the force on the cyclic stick in flight using a hand held force gage or other force measuring instrumentation. The in- flight technique is the same as the first method. Testing should be accomplished at a minimum of two altitudes. One altitude should be low enough to assure limiting power is attained.
Another should be at or near the maximum approved altitude. Reasonable interpolation is allowed. If no marginal areas are apparent interpolation over a 10,000-foot altitude range is considered reasonable.
(ii) Tests for static longitudinal stability during approach should include the steepest approach gradient for which approval is requested. Static stability tests may be simulated by initially establishing a trimmed rate of descent for maximum approach gradient assuming zero wind conditions. Actual approach tests at the maximum approved gradient should be conducted to evaluate tracking and maneuverability, including the capability to correct downward to a glide path when approaching in a slight (10 knot) tailwind condition.
(4) Static Lateral Directional Stability.
(i) Tests for directional stability usually require instrumentation for lateral cyclic position, pedal position, and sideslip angle. Testing for compliance with the specific directional requirement is relatively simple; however, the pilot should look for significant longitudinal trim changes and short period dynamic modes, which might occur only during sideslip conditions. Side force characteristics are indicated by the variation of bank angle with sideslip during steady heading sideslips. The number of ball widths of deflection is also indicative of the side force cue available to the pilot. A correlation between sideslip angle and ball widths of skid can be obtained at given speeds for use during later testing after sideslip instrumentation is removed. A simple yaw string can be calibrated in a similar manner. The TIA should define the maximum sideslip angles, which should not be exceeded during the flight test program. These angles must not be greater than the structural sideslip envelope substantiated and are not required to be that sideslip angle obtained with full directional pedal deflection.
Sufficient side force cues should accompany sideslip to alert the crew when approaching sideslip limits. This is needed to assure that structural sideslip limits will not be inadvertently exceeded in service. Although not stated in the requirement, flight conditions for demonstration of static longitudinal stability are also appropriate for demonstration of static lateral-directional stability.
(ii) Dihedral requirements may be more difficult to assess. For those rotorcraft that do not meet the position and force gradient requirements for the conventional, cross-controlled sideslips, there are alternative tests that may be used to determine acceptable characteristics. If directional pedals are util ized in steady sideslips, the resultant rolling tendency is the sum of (1) the aircraft’s roll due to sideslip tendency (dihedral) and (2) the aircraft’s roll due to directional control input. If the rotorcraft has a tail rotor, which is excessively high or low in relation to the rotorcraft’s vertical center of gravity (CG), application of tail rotor thrust will introduce a significant Page Apdx B - 3
AC 29-2C, Chg 7 2/4/2016 rolling moment. The basic intent of dihedral stability testing is to determine the rotorcraft response to sideslip exclusive of directional control input. In general, if a tail rotor configuration is involved and the tail rotor is above the vertical CG of the rotorcraft, the effect of pedal input upon dihedral effect is destabilizing during conventional, control-induced sideslips.
(iii) There are two alternate methods, which for small angles of sideslip, can give an indication of the basic dihedral stability of the rotorcraft. Both methods involve freezing directional controls while artificially creating sideslip by other means.
(iv) The first method is only applicable for rotorcraft with single main rotor systems. To utilize this method, the rotorcraft is stabilized in a given flight condition and small collective (torque) changes are applied in each direction (e.g., ±5% and ±10%) while holding pedals fixed. Sideslip angle, lateral control position, and lateral control force may be measured and plotted for small torque changes from trim. This technique will not work for aircraft that have collective to pedal or collective to lateral control couplings.
(v) In the second method, the rotorcraft is stabilized in a trimmed flight condition with a small amount of bank (5-10°). The rotorcraft is then rolled to an approximately equal angle of bank in the opposite direction holding the pedals fixed.
The change in direction of bank results in a small change in sideslip angle and again sideslip angle may be plotted versus lateral control position or force. This test should be conducted in both directions and the results averaged. This method can give reasonably accurate results for small perturbations. Other factors contribute to the results of either of these two methods. It is always important to assess the roll due to sideslip tendency with pedal induced sideslips to assure lateral control forces are reasonable and in a proper direction for directional out-of-trim conditions, and to assure the pilot has adequate sideslip cues.
(vi) Wording of the dihedral requirement is intended to allow slightly negative dihedral stability at critical loading conditions. This will ordinarily result in positive dihedral stability throughout a great majority of the approved loading envelope. The test for maximum allowable negative dihedral effect would involve stabilization at a required flight condition, inducing a sideslip up to ±10° from trim, then assessing lateral cyclic friction or deadband to determine if roll is restrained while remaining in the control system friction or deadband so that the control may be released without resulting in the aircraft rolling in the adverse direction. When testing for this condition, lateral cyclic friction should be adjusted to the minimum value.
(vii) The intent of the dihedral rule is to allow small amounts of control system friction and deadband to mask small values of negative dihedral. Where slope of the negative dihedral versus sideslip exceeds these small values, the negative dihedral must not be approved. The operational pilot must not be presented with opposite cyclic sensing for similar sideslip conditions as loadings and flight conditions change. In general, large values of control system friction and deadband are undesirable. The Page Apdx B - 4
addition of friction or deadband into the control system for satisfying the dihedral requirement is not acceptable.
(viii) In approving small negative dihedral values, the pilot should ensure that other positive flight cues, such as suitable side force, accompany sideslip. This will aid the pilot in determining direction of sideslip so that no reverse sensing or confusion accompanies sideslip conditions.
(5) Dynamic Stability.
(i) Dynamic characteristics are defined in quantitative terms; however, some areas of interpretation and technique need special consideration:
(A) Unlike fixed-wing aircraft where the size of the input has no effect on damping ratio, rotorcraft can be sensitive to the type and size of input used to excite each dynamic mode. For instance, it has been found that for the phugoid-type dynamic oscillation, damping ratio is inversely proportional to the size of the input. It therefore becomes important that dynamic excitations be sized to approximate the response of the rotorcraft in a moderate turbulent gust. In addition, the dynamic input should be made with the control(s) that most accurately simulates the typical aircraft gust response. Obviously, for this evaluation some flying of the rotorcraft in turbulence is necessary to obtain knowledge of the rotorcraft’s gust response. Pulses and doublets may be used to generate disturbances similar to a gust. To assist returning the control(s) to the trim position a hand held jig may be used. Use of attitude and rate instrumentation is desirable. The pilot may find that collective excitation, or collective in conjunction with cyclic, is most appropriate for gust simulation.
(B) The second area of concern in evaluating dynamic response is whether to let only one axis respond to an excitation or to let the rotorcraft respond in two or more axes. When it can be done safely, the rotorcraft should be allowed to follow its dynamic response in all axes. In other words, if pitch oscillations feed into roll, the pilot should attempt to observe and record the total aircraft dynamic response in both pitch and roll.
(C) The third area concerns strict compliance with the exact wording of the dynamic requirement. In this regard, a neutrally damped oscillation with a period of 19 seconds would not be acceptable; however, a very divergent oscillation that doubles in amplitude in 21 seconds would be acceptable. The 19- second oscillation is much less severe than the 21-second oscillation and yet is unacceptable by the “letter of the law.” Figure AC 29 APX B-1 is a graphical display of the dynamic requirement. The 19 and 21-second oscillations are shown as points (1) and (2). Point No. 1 is positioned much more toward the acceptable portion of the graph and yet by the “letter of the law” is unacceptable. The intent of the dynamic requirement is roughly approximated by the dashed or curved line. Areas to the right of that line may be considered for findings of equivalent safety.
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AC 29-2C, Chg 7 2/4/2016 (D) A fourth area requiring special care in testing is the aperiodic requirement. The most common aperiodic motion is the spiral characteristic, which results when aircraft attitude is displaced in roll. The preferred method for testing this requirement is to stabilize precisely on a trimmed condition in straight flight, then displace the rotorcraft to 10° of bank, stabilize momentarily, set the controls as they were positioned for straight flight, and release them. Time and bank angles are then recorded. Recovery is initiated when bank angle or roll rate becomes excessive. Of particular interest is the time for bank angle to pass 20° and this time should not be so short as to cause the aircraft to have objectionable flight characteristics in the IFR environment. The time period to double amplitude (20°) should be at least 9-seconds.
It is vitally important that controls (particularly lateral cyclic) are positioned exactly as it was for the straight flight condition. If a high resolution force trim system is not incorporated, an alternative method may be used. In this second method, the rotorcraft is trimmed for straight flight as described above and controls are released. Roll attitude may simply be allowed to vary naturally with time or small pulse input may be made with pedals. It is important that controls be positioned precisely as they were for the trimmed, straight flight condition and a plot of bank angle versus time is obtained. This plot is then compared against a divergent roll condition that doubles in amplitude every 9-seconds. Of particular interest is again the rate passing 20° of bank. If airspeed changes as the aircraft rolls or if roll or pitch coupling occurs, these changes should be allowed to interact naturally until recovery is necessary. Due to the sensitive nature of this test, smooth air is essential. Repeatability may be a problem. At least two test points in each direction should be obtained at each trim condition. Results may be averaged if they show reasonable repeatability. The same procedures may be utilized for an aperiodic pitch response; however, a displacement of 5° from trim should be used and of particular importance is the pitch rate passing 10°. Again, at least two test points in each direction should be obtained for each trim condition. Although not stated in the requirement, the flight conditions for demonstration of static longitudinal stability are also appropriate for demonstration of dynamic stability. The degree of testing referred to here represents that which might be required of a marginally stable rotorcraft. For those configurations that provide good aerodynamic stability or use varying degrees of SAS, the scope of the demonstration program would be decreased significantly.
(ii) Control system dynamics should also be evaluated. This may be accomplished by lightly bumping each control in flight and observing its free response.
Any resulting control motion must dampen quickly and should not be driven by aircraft or control system interaction. This will assure safe flight in the event a control is inadvertently bumped or released from an out-of-trim condition.
(6) Automatic Flight Control System (AFCS).
(i) From a regulatory compliance standpoint, the AFCS used on most modern rotorcraft often performs two different and distinct functions. These two functions are to:
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(A) enhance the inherent basic stability of the aircraft to enable compliance with applicable flight characteristic requirements by using automatic control features such as SAS, which may incorporate an attitude hold mode (ATT); and (B) control aircraft flight path using a simple ATT in conjunction with the SAS or autopilot modes (coupled flight director), such as airspeed hold, altitude hold, radio navigation, and precision approach.
(ii) Definitions.
(A) Autopilot (AP). The AP, also known as coupled flight director, is an automatic mechanical, electrical, or hydraulic control system primarily intended to assist the flight crew in the guidance of the aircraft. The various AP control modes are designed to provide workload relief to the pilots and a means to fly a flight path more accurately to support specific operational requirements.
(B) Attitude Hold Mode (ATT). ATT is an automatic flight control system mode that maintains established pitch and roll attitudes and, in some cases, provides yaw damping, yaw hold, or heading hold.
(C) Flight Director (FD). The FD is a guidance system that provides an automatically computed visual cue or set of cues, used during manual control of the aircraft as steering command information, which direct the pilot how to maneuver the aircraft (usually in pitch, roll, and collective) to track a desired flight path.
(D) Stability Augmentation System (SAS). The SAS is an automatic control system mode that augments the inherent stability of the aircraft. The SAS aspects of the AFCS are not covered in § 29.1329 (Automatic pilot system) but rather in § 29.672. Nevertheless, the guidance provided in this section for pilot response times is applicable to SAS malfunction testing.
(iii) AFCS Malfunctions.
(A) AFCS Malfunction Assessment. Failure conditions in the AFCS, including SAS, ATT, FD, and AP functions, that are not shown to be extremely improbable must be demonstrated in flight. Assessment of failure conditions has the following elements:
(1) Failure condition insertion.
(2) Pilot recognition of the effects of the failure condition.
(3) Pilot reaction time (that is, the time between pilot recognition of the failure condition and initiation of the recovery).
(4) Pilot recovery.
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AC 29-2C, Chg 7 2/4/2016 (B) Failure Condition.
(1) Failure Conditions. The AFCS failure conditions should be simulated such that overall response is representative of the actual effect of that failure condition on the aircraft and its systems.
(2) Most Critical Conditions. The aircraft loading and flight conditions under which the failure condition is inserted should be the most critical. Examples of these conditions are CG, weight, altitude, speed, and power. If the AFCS includes coupling to the collective axis, the tests should be performed with the collective axis both coupled and uncoupled.
(C) Pilot Recognition and Response Time.
(1) The pilot recognition and response time is dependent on pilot involvement in the flying task, which is defined as “active,” “attentive,” and “passive” flight. The pilot’s involvement in the flying task is dependent on the level of stabilization provided by a particular mode of the AFCS, as well as the phase of flight.
(i) Active flight - any flight segment during which the characteristics of the aircraft necessitate continuous flying of the rotorcraft by the pilot using the primary flight controls. Active flight includes take-off, hover, and landing.
(ii) Attentive flight - any flight segment requiring attention from the pilot for short periods. Attentive flight can be further subdivided into hands -on and hands-off. Attentive, hands-on addresses phases of flight where the AFCS is controlling the flight path, but close attention is required from the pilot. Examples include coupl ed hover and coupled approach. Attentive, hands -off addresses flight phases where the controls can be released for extended periods of time, but where the pilot is expected or required to continue to monitor the behavior of the aircraft and system. Examples include using the following AFCS modes in climb, cruise, hold, and descent: a more basic attitude hold such that the pilot is required to monitor the series actuator authority and correct as necessary; or an AP mode if an AFCS malfunction results in a caution message and the applicable emergency or abnormal procedure requires the pilot to be attentive.
(iii) Passive flight - any flight segment of long duration requiring the minimum of attention from the pilot. This is intended to address the use of AP modes of the AFCS in climb, cruise, hold, and descent. It is also applicable to attitude retention mode where the characteristics of the system do not require continuous close monitoring of system behavior.
Note: It is not acceptable to use reduced pilot response times based on rotorcraft flight manual (RFM) limitations requiring the pilot to be hands-on if hands-on flight is not appropriate to the level of stabilization and flight path control provided by the Page Apdx B - 8
AFCS operating mode.
(2) Pilot recognition of AFCS malfunctions. The pilot may detect an AFCS failure condition through aircraft motion cues or by cockpit flight instruments or alerts. The specific recognition cues will vary with specific failure type, flight condition, phase of flight, and crew duties. When evaluating AFCS malfunctions for attentive hands-off and passive flight, it should be assumed that the flight crew is performing cockpit duties that require it to divert their attention away from the flight instruments.
(i) Hardover. The recognition point should be that at which a pilot operating in non-visual conditions may be expected to recognize the need to take action. Recognition of the effect of the failure may be through the behavior of the rotorcraft or through an appropriate alerting system. Flight control movements alone should not be used for recognition. The pilot recognition time will be the least of the time taken for the:
(A) rotorcraft to achieve an angular rate of change about any axis of 3° per sec (active and attentive flight). The time taken for the rotorcraft to achieve an angular rate of change about any axis of 5° per sec (passive flight);
(B) rotorcraft to increase or decrease acceleration along any axis by 0.2g (active and attentive flight). The time taken for the aircraft to achieve an acceleration about any axis of 0.25g (passive flight); or (C) relevant alert to function.
(ii) Slowover. This type of failure condition is typically recognized by a path deviation indicated on primary flight instruments (for example, course deviation indicator (CDI), altimeter, or vertical speed indicator). It is important that the recognition criteria are agreed upon with the FAA. The following identify examples of recognition criteria as a function of flight phase:
(A) En-route cruise. Recognition through the altitude alerting system can be assumed for vertical path deviation. The lateral motion of the rotorcraft may go unrecognized for a significant period of time, unless an alerting system is installed and the slowover results in a path or attitude deviation in excess of the limit allowed by the alerting system.
(B) Climb and descent. Recognition through increasing or decreasing vertical speed, pitch attitude, roll attitude, or heading can be assumed.
(C) On an approach with vertical path reference. A displacement recognition threshold should be identified and selected for testing that is appropriate for the display(s) and failure condition(s) to be assessed.
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AC 29-2C, Chg 7 2/4/2016 (D) On an approach without vertical path reference. Criteria similar to the climb or descent condition can be assumed.
Note: Credit may be taken for excessive deviation alerts, or other clear and unambiguous alerts provided to the pilots, if available.
(iii) Oscillatory. Oscillatory failures having structural implications must be addressed under part 29, Subpart C and § 29.1309. The flight crew is expected to disengage the automatic control elements of the AFCS that have adverse oscillatory effects and not follow adverse oscillatory guidance. However, if unable to disconnect some elements of the AFCS in the presence of an oscillatory failure condition, there should be an evaluation of the long-term effects on crew workload and the occupants.
(3) Pilot response time. Pilot response time consists of pilot recognition time + pilot reaction time. The pilot response time is considered dependent upon the pilot attentiveness, based upon the phase of flight and associated duties. This period commences at the time the pilot is cued to the fact that something abnormal is occurring (at the end of the pilot recognition time) and terminates when the controls are moved to commence the recovery maneuver. The pilot recognition time is assumed to increase as the pilot relaxes his involvement level. The pilot reaction time is longer for 'hands-off’ than for 'hands-on' as the pilot has to locate the controls before he can move them. The following pilot response times have been considered acceptable:
Flight Segment Recognition Time (sec) Reaction Time (sec) Pilot Response Time (sec) Active - 0.5 0.5 Attentive–Hands-On 1.0 0.5 1.5 Attentive–Hands-Off 1.0 1.0 2.0 Passive 2.0 1.0 3.0 Notes: Recovery from simulated malfunctions of any AFCS axis occurring while the pilot is applying control inputs to cause rotation about that axis may be initiated with normal pilot reaction; the 1-second delay (decision time) in maneuvering (attentive hands-on) flight pertains to established turns (level, climbing, and descending) only.
For rotorcraft requiring a minimum crew of two pilots and with AFCS systems that do not have the capability to couple to AP modes, it can be assumed that the pilot is attentive-hands-off during the climb, cruise, and descent phases of flight.
(D) Pilot Recovery.
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