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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(i) Pilot recovery action. The recovery action should be commenced after the pilot recognition and response times. Following such delay, the pilot should be able to return the rotorcraft to its normal flight attitude under full manual control without excessive pilot skill or strength. During the recovery, the pilot may overpower the autopilot or disengage it.
(ii) AFCS Altitude Loss Testing. To determine the maximum altitude loss of the various AFCS modes during the different flight phases, a representative pilot recovery maneuver appropriate to the rotorcraft type and flight condition should be performed. This maneuver should not lead to an unsafe speed or attitude excursions to resume a normal flight path. An incremental normal acceleration in the order of 0.5 g is considered the maximum for this type of maneuver.
(7) Stability Augmentation System (SAS).
(i) If an SAS installation stabilizes the rotorcraft by allowing the pilot to “fly through” and perceive a stable, well-behaved vehicle, it qualifies as a SAS, and if reliable, receives credit under paragraph III. through VII. of Appendix B to part 29 for use in complying with all-handling qualities requirements. If a conventional autopilot does not provide “fly through” capability or allow the pilot to perceive a stable, well behaved vehicle through his manipulation of primary flight controls and feedback from those controls, then it tends to remove him from active involvement in flying and is eligible primarily as a workload reliever.
(ii) If handling qualities credit is given for a SAS then it must be shown to be reliable in regards to § 29.1309. If a reliable SAS is incorporated, it should be operational during handling qualities testing for trim and stability. Reasonable single failures of the SAS must be evaluated and the resultant handling qualities must be evaluated to assure that in this degraded configuration: (1) handling qualities have not been degraded below “VFR” levels defined in part 29, Subpart B; (2) the rotorcraft is free from any tendency to diverge rapidly from stabilized flight conditions; and (3) the rotorcraft can be flown IFR throughout its endurance capability without undue difficulty by the minimum flight crew. Compliance with a majority of the IFR handling qualities requirements is desired and the degraded characteristics should be documented and explained. Revised flight envelope boundaries for the failed condition may be considered if they are controllable by the pilot (e.g., altitude and airspeed). When loss of a SAS results in a need for minor adjustment of a flight condition then a system can be accepted that allows failures during the life of each rotorcraft. If loss of the system will prevent continuation of safe flight and landing, the reliability of the system must be high enough to assure that failure of the system will not be expected to occur during the life of the rotorcraft fleet. When evaluating the reliability of a system, the installation of the system should be considered as part of the design. The total system including inputs, outputs, environment, isolation features, and exposure times is a pertinent consideration.
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- AC 29-2C, Chg 7 2/4/2016
(iii) SAS reliability is evaluated by Systems and Equipment personnel. If credit is to be given for system reliability and the applicant exempted from consideration of malfunction, hardover and oscillatory conditions (limited to critical frequencies determined during autopilot failure analysis), a thorough system evaluation is needed.
Flight test personnel should coordinate closely with the systems and equipment personnel whenever credit is given for advanced design and system reliability because the hardover or malfunction condition may not require in-flight testing. The decision is made on the basis of system design, failure analysis, and overall probability of malfunction. Multi-lane systems require careful consideration when determining the need for hardover or malfunction testing of second and subsequent failures. Factors to be taken into account are:
(A) Dissimilarity of architecture (hardware and software) between the multiple lanes.
(B) Cumulative probabilities. For example, in a quadruplex system where each lane has a failure probability of 1x10-4 and depending on system architecture, the level of criticality applied to the software, and the potential for generic failures the probability of three lanes failing leaving the remaining lane unmonitored would be a 1x10-12 event. In this case, testing of the third and fourth failures would not be required.
(C) Consequences of a multiple failure, considering § 29.1309. A multiple failure should not present a hazard to a category B rotorcraft or prevent continued safe flight and landing for a category A rotorcraft.
(D) Operational restrictions post-failure. If it were possible to recouple the upper modes and continue the flight without restrictions post-failure, then consider any subsequent failures with pilot intervention times relevant to the phase of flight. If a flight restriction is provided, it should be determined to be an appropriate and relevant operating limitation, and it should be specified in the RFM. Significant information regarding the restriction should be made available to the pilot in the operating procedures section of the RFM. If the restriction excludes operation under any of the flight conditions listed above, flight testing of the condition is not required. If the system is to be approved without flight restrictions (operating at all times), malfunctions should be demonstrated to be satisfactory during takeoff, climb, cruising, landing, maneuvering, and hovering.
(iv) Applicable procedures and techniques for conduct of hardover tests are contained in section 29.1329 of this AC. All cockpit emergency controls including emergency quick disconnects should be “red.” The quick disconnect may be actuated at initiation of recovery. Other disconnects should only be actuated after full aircraft control has been achieved following recovery. Aircraft limits may not be exceeded during malfunction or recovery. If a monitor device automatically disconnects the SAS, it must be clearly annunciated to the crew. Appropriate delay times are shown in paragraph b.(6)(iii)(C)(3).
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(v) Series actuator hardover conditions in some rotorcraft can seriously degrade control margin. Critical loadings, power settings, RPM, and altitudes in conjunction with a SAS actuator hardover in an adverse direction can result in reduction of control travel requiring flight envelope constraints. Flight testing is usually necessary to determine the appropriate flight envelope reductions.
(vi) Subsequent failures and unrelated probable combinations of failures must be considered, including subsequent SAS failures. Systems and equipment section analysis should provide necessary SAS malfunction combinations for flight testing because of their system analysis. Minimum requirements for dispatch and procedures following failure should be included in the malfunction analysis. Results of the probability analysis and the resultant malfunction configurations are primarily the responsibility of the systems and equipment section.
(vii) No reasonably probable failure should result in a worse condition than that tested for hardovers. For example, if a magnetic brake force trim system is employed, failure of electrical power to the magnetic brake circuit may cause the cyclic control to fall, which may result in a more dangerous flight condition than individual SAS hardovers. The overall control system is to be evaluated for all probable failures to preclude hazardous failure conditions. Other areas for investigation include beep trim and auto trim failures. The delay times of paragraph b.(6)(iii)(C)( 3) are appropriate for all such failures. System malfunctions may also include component failures, which result in oscillatory outputs of the actuator(s). These should be sustainable at least as long as the specified hardover delays, should be manageable thereafter with hands on the controls, and should allow disconnect of the malfunctioning system.
(viii) Engine failure requirements are not entirely consistent with the SAS failure time delays shown in paragraph b.(6)(iii)(C)(3). Engine failure time delays remain as specified in § 29.143(e) and they are lower than corresponding SAS failure delays. Critical engine failure conditions should be reverified during simulated instrument flight with primary reference to flight instruments. Lower time delays for engine failure have been justified on the basis of immediate cues for the critical high powered condition, and requirements for engine failure warning systems. Many rotorcraft designs simply cannot endure a 3-second time delay for critical engine failure conditions. Nevertheless, engine failure, autorotation entries, and autorotation descent (for single engine rotorcraft and multiengine rotorcraft without category A engine isolation) must be evaluated in simulated IFR conditions and these flight characteristics must be acceptable.
(8) Controllability.
(i) Control harmony should be present. There should be no objectionable cyclic to collective or roll-yaw-pitch cross coupling.
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AC 29-2C, Chg 7 2/4/2016
(ii) Control forces following a control system malfunction such as a hydraulic system failure should be low enough to allow completion of the intended flight. It may not be possible to land early during an actual IFR flight.
(iii) There should be no tendencies for pilot induced oscillations; there should be no sustained or uncontrollable oscillations resulting from the efforts of the pilot to control the rotorcraft.
(iv) The control system must have sufficient resolution to permit accurate and precise instrument maneuvers. Some control systems with high breakout forces in conjunction with low control force gradients do not lend themselves to satisfactory instrument flight capability.
(9) Cockpit Arrangement.
(i) The primary flight instrument basic T (or a modified T with VSI above the altimeter) should be located directly in front of the pilot. All annunciations necessary for operation of stability systems should be readily in view. Secondary flight (or navigation) instruments such as radar altimeter and secondary radio course information, DME, etc., should be grouped around the periphery of the T. Next in priority are primary power instruments such as torque and rotor RPM. Powerplant instruments and backup attitude information should be placed in the remaining panel areas. Various research and development efforts and previous certification programs have revealed that it is desirable not to locate the standby attitude indicator immediately adjacent to the basic flight instrument T. The standby attitude indicator must be usable and flyable from the primary pilot station (and any other pilot station); however, locating it too close to the primary instruments may be undesirable and should be evaluated. If the standby attitude information is close to the pilot’s normal flight instrument scan, he may begin to compare attitude information between the two indicators in his normal instrument scan.
Every pilot eye motion to compare these indicators could be a wasted motion that could be more efficiently applied in the normal scan. The pilot should fly either the primary or the backup indicator and it may be an aid if these indicators are noticeably separated.
When the standby indicator is located apart from the normal scan and the primary indicator fails, the pilot is conscious of a distinctly different instrument scan and is less likely to be continuously coming back to the center of the basic T for attitude reference.
Physical separation can assist the transition to standby attitude flight.
(ii) All cockpit controls necessary for normal and emergency operations should ideally be located so that they may be actuated without upper body movement.
Moderate head and body movement has been accepted; however, these motions must be evaluated for their vertigo inducing effects. No IFR controls should be located aft of a vertical plane passing left to right (laterally) through the pilot’s body.
(iii) If a copilot position is approved, the copilot must have a complete set of flight controls, and a complete set of primary flight instruments. The copilot must be capable of independently flying and navigating the rotorcraft from his position. The Page Apdx B - 14
copilot must be capable of controlling at least one primary navigation source so that he can operate the rotorcraft during normal conditions without relying on the first pilot to perform needed cockpit functions. Some instruments can be shared between pilots depending on instrument panel presentation. Some examples from previous programs include standby attitude, rotor tachometer (if the aircraft has automatic governing and the crew is provided visual and aural RPM warning), and secondary powerplant instruments such as NG, oil pressure, and temperature.
(iv) Proper cockpit annunciation is essential for safe operation. SAS and autopilot modes must be properly annunciated. Appropriate annunciator color coding is contained in § 29.1322. There must be no question in regard to the source of navigation information presented to the crew. Where navigation switching is available between individual displays and between pilot positions, the first pilot should have overriding control for his displays.
(v) Electromechanical Displays. Due to the increased complexity of instrumentation that is available and in use, it is appropriate to consider the extreme range of operational environments to which rotorcraft were being routinely exposed. It is the intent of Appendix B to part 29, VIII.(b)(5) to prevent degrading of the first pilot’s instrument system, or the only pilot’s instrument system in a single-pilot-approved rotorcraft, by not permitting connection of peripheral systems to it. In addition, equipment must not be connected to operating systems for the second pilot’s required instruments unless it is extremely improbable that failure of such additional equipment would affect that operating system.
(vi) Advanced Display Systems. The increased use of mi croprocessor technology in avionic systems has resulted in the use of computer-generated graphics to replace conventional electromechanical instruments. These displays may replace individual instruments or may integrate several flight critical parameters into single displays. For display of redundant information, “crosstalk” between the pilot and copilot displays and supporting systems has been allowed to provide detection and annunciation of faults or “miscompare” of critical flight information. A level of safety finding equivalent to that level of safety provided by Appendix B to part 29, VIII.(b)(5)(i) through (iii) may be possible through the implementation of integration technology that will assure that failure of one system does not and can not adversely affect the other system. For those installation designs that employ integration technology, adequate system testing and any analysis necessary must be conducted to assure that failure of one system will not adversely affect the other system when demonstrating compliance to the minimum safety level established by Appendix B to part 29, VIII.(b)(5)(i) through
(iii).
(10) IMC Evaluation.
(i) As part of the flight test program, new rotorcraft undergoing IFR certification should be flown in the air traffic control system in actual day and night Page Apdx B - 15
AC 29-2C, Chg 7 2/4/2016 instrument meteorological conditions. Items for consideration during the IMC evaluation include:
(A) Ability of the rotorcraft to safely operate in the National Airspace System, including crew capabilities to cope with probable malfunctions. Examples of failures imposed during this IMC evaluation on previous programs are shown below:
(1) Hydraulic failure.
(2) Individual COMM, NAV, or intercom failure.
(3) Engine failure.
(4) Loss of any power input.
(5) SAS failure.
(6) Trim failure.
(7) Individual failure of each vertical and directional gyro.
(B) Visibility during low approach conditions in precipitation.
(C) Glare and reflections at night in clouds.
(D) Workload demands on the minimum flight crew including the failures in paragraph b.(10)(i)(A).
(E) Handling qualities in turbulence throughout the IFR approved envelope including typical IFR flight maneuvers with:
(1) reasonably anticipated SAS failures,
(2) reasonably probable control system failures (hydraulics, force trim, basic ship systems, etc.),
(3) the typical workload conditions associated with operating in high density traffic areas, and
(4) other reasonable, probable failures.
(F) Cockpit leaks in precipitation that affects pilot efficiency, safety, or rotorcraft airworthiness.
(ii) Rotorcraft that are an improved, modified, or later model of previously approved type that have no significant changes in the fuselage and windshield Page Apdx B - 16