from CAR Part 7. However, § 29.663 was adopted by Amendment 29-3 to investigate possible sources of ground resonance and to assure that the reliability of the ground resonance prevention means; i.e., dampers, shock struts, etc., would preclude the occurrence of ground resonance. The total rotorcraft system, including landing gear, struts, tires, etc., is evaluated under this standard.
(4) Viscous dampers in the rotor head have been used for many years to prevent ground resonance. Modern rotorcraft designs may also use elastomeric dampers and may use elastomeric bearings in the rotor head and rotor pylon attachment to the airtrame. The standard applies to viscous and elastomeric dampers.
The "probable" range in damping shall be investigated. The standard also requires investigation of the probable range of variations of these dampers, whether viscous or elastomeric, and elastomeric bearings to preclude ground resonance.
(5) Ground resonance can occur due to flexibility in the rotor pylon restraint system as well as with landing gear flexibilities and/or shock struts. See Paragraph b(2) for an explanation. An analysis may be done to show the effect of the rotor pylon mount stiffness on ground resonance stability. If the analysis shows that rotor pylon mount stiffness could affect ground resonance, the evaluation should include variations in stiffness and damping of the rotor pylon restraints that may occur in service (reference "Ground Vibrations of Helicopters," M.L.. Deutsch, JAS, Vol. 13, No. 5, May 1946).
b. Procedures.
(1) The reliability of the means for preventing ground resonance may be substantiated as stated in the standard. An analysis report or a test proposal and subsequent test report may be used to show compliance. The probable range of variations, in service, of the damping action are an important part of the assessment.
The test may be conducted in conjunction with the testing required by § 29.241. See Paragraph 99.
Analysis and tests may be used.
(її) Reliable service history of identical or closely similar systems may be used. The materials and fluids used, clearance or fits, seals, and physical installation are important items to be evaluated and considered for "closely similar" systems.
(Ш) Testing of the complete rotorcraft may be used to prove that malfunction of a single means of the damping system will not cause ground resonance.
One method of demonstrating acceptable compliance is by removing all seals, if practicable, from one damper. Another method is to remove all or most of the fluid, in conjunction with considering the allowable ranges of damping of the other parts of the rotorcraft damping system and operating the rotorcraft throughout the rotor speed range from start to maximum rotor speed. Investigation of elastomeric dampers may require 402 Par 268
innovative test procedures and preliminary discussions of these prior to preparation of a test proposal. The rotorcraft cyclic control should be displaced as noted in Paragraph 99 of this document to assure that the possible rotorcraft resonance frequencies are excited. If vibrations are damped in all tests, the damping system is satisfactory. Each critical rotor damper and landing gear damper, which includes shock struts and tires, should simulate a malfunction to comply with the standard. The testing discussed, however, could become very extensive if one were to attempt to test all combinations of all maintenance adjustments of all components which contribute to the prevention of ground resonance, while at the same time rendering each of the pertinent components ineffective in turn and then repeating all of the maintenance tolerance testing each time. Fortunately, rational analytical methods are available which will permit the evaluation of such combinations so that only the combinations with the least amount of margin used are physically tested.
(2) The pylon damper variation can affect ground resonance. The variations in stiffness and/or damping of pylon mounts should be evaluated except the pylon mounts on contemporary conventional rotorcraft may have little influence on "classical" ground resonance stability. The dynamics of the rotorcraft on its landing gear is generally established by the airframe properties and the landing gear properties under the influence of the rotor system, with the "pylon" having little effect. For air or flight resonance, the rotor generally couples with the rigid body modes of the fuselage. For a specific design, a relatively simple analysis may be used to show the effect of the pylon mount system stiffness on air and ground resonance stability, and if not important, variations in the system may be omitted from the test program.
(3) The probable ranges of damping shall be established and investigated as prescribed and noted in Paragraph (b) of § 29.633. An approved test proposal and test results report should be used for complying with § 29.663(b). For example, if a conventional wheel landing gear is used on the rotorcraft, the probable ranges of tire pressure or the lowest probable tire pressure should be stated in the test proposal and effects of the tire pressure investigated during the test. In addition, the effects of strut pressures should be investigated also. See Paragraph 99, § 29.241, concerning tests and instrumentation of the test associated with complying with § 29.241. The instrumentation noted in Paragraph 99 also applies to § 29.663(b).
(4) If the wheel landing gear is equipped with wheel brakes, the evaluation should include brakes "on" and "off." The nose or tail wheel should be locked and unlocked if it swivels to evaluate any possible adverse effects of this feature.
268A. § 29.663 (Amendment 29-30) GROUND RESONANCE PREVENTION MEANS
a. Explanation. Amendment 29-30 clarifies that analysis as well as tests may be used to show freedom from ground resonance after malfunction or failure of a single Par 268 403
means of ground resonance prevention. This amendment primarily clarifies that the probable range of damping should be established as well as investigated.
b. Procedures. The procedures of Paragraph 268 continue to apply with the addition of the need to document the establishment of probable range of damping of ground resonance prevention means. Acceptable tire and oleo minimum and maximum pressures as well as other identified factors should be documented in maintenance instructions if necessary to assure the desired characteristics.
404 Par 268A
269.-275 RESERVED.
Intentionally Left Blank Par 269 405 (thru 411)
SECTION 17. CONTROL SYSTEMS
276. § 29.671 (Amendment 29-24) GENERAL.
a. Explanation.
(1) The rule requires that controls operate easily and smoothly and provide positive response of the rotorcraft to control input.
(2) in addition, the rule requires that incorrect assembly be prevented by special design features or special markings.
(3) After Amendment 29-24, November 6, 1984, the rule requires that the flight control system be designed such that the full range of flight control authority can be verified by the pilot before flight. This check would normally have to be completed prior to turning the rotor since control extremes typically cannot be reached with the rotor operating on the ground.
b. Procedures.
(1) Easy, smooth operations of controls are substantiated by the operations tests of § 29.683 and the FAA/AUTHORITY flight testing under TIA procedures.
Positive response of the rotorcraft to control inputs is also evaluated during company flight testing and FAA/AUTHORITY TIA flight testing to the requirements of §§ 29.141 through 29.175.
(2) To meet the requirement that incorrect assembly be prevented, the preferred method is providing design features which make incorrect assembly impossible. Typical design features which can be used are different lug thicknesses, different member lengths, or significantly different configurations for each system component. In the event that incorrect assembly is physically possible (because of other considerations), the rule may be met by the use of permanent, obvious, and simple markings. Permanent (durable) decals or stencils may be used.
(3) Design features of the control systems are checked when reviewing the type design drawings. During the proof and operation tests of §§ 29.681 and 29.683, the controls should be thoroughly reviewed for possible incorrect assembly and for any required markings supplied for compliance with this standard.
277. § 29.672 (Amendment 29-24) STABILITY AUGMENTATION, AUTOMATIC.
AND POWER-OPERATED SYSTEMS.
a. Explanation.
412 Par 276
(1) This rule requires that the pilot be made aware of stability augmentation, automatic, or power-operated system failures which could lead to an unsafe condition.
Examples of clearly distinguishable warnings include, but are not limited to, an obvious aircraft attitude change following the failure or an audio warning tone. A visual indication itself may not be adequate since detection of a visual warning would normally require special pilot attention. The use of devices such as stick pushers or shakers is not acceptable as a warning means. However, this rule is not intended to eliminate the use of such devices for other purposes. Examples of automatic control systems other than a stability augmentation system would be a pitch axis actuator used for the purpose of demonstrating compliance with longitudinal static stability requirements or a fly-by-wire elevator. The design of such systems must not interfere with completion of the control checks described in § 29.671 (c). Further, for control systems where a series actuator malfunction could degrade control authority, a means should be provided to the pilot to determine actuator alignment (see § 29.1329(b)).
(2) The corrective flight control input following a system failure should be in the logical direction. For example, a malfunction resulting in a nosedown pitch of the aircraft should require a corrective cyclic control input in the aft direction. The system deactivating means does not have to be located on the primary flight control grips;
however, it should be easily accessible to the pilot. Malfunctions and subsequent recoveries must be shown throughout the operating envelope of the aircraft. In a case where control authority is decreased following a malfunction, a reasonable flight envelope must be defined wherein compliance with controllability and maneuverability requirements can be demonstrated. This reduced flight envelope must be presented in the flight manual. Compliance with trim and stability characteristics is not required following a malfunction; however, a pilot workload assessment should be made to show that a mission can be safely continued to completion following the worst case single failure.
b. Procedures. A discussion of malfunction test procedures is presented in Paragraph 775b(6). Controllability and maneuverability test procedures are addressed in Paragraph 80.
278. § 29.673 (Amendment 29-24) PRIMARY FLIGHT CONTROLS.
a. Explanation. This section basically defines primary flight controls as "those used by the pilot for immediate control of pitch, roll, yaw, and vertical motion of the rotorcraft. This section clarifies the application of § 29.1555 which requires markings for controls other than "primary flight controls or controls) whose function is obvious."
b. Procedures. The primary flight controls (e.g., cyclic stick, collective, and tail rotor pitch control pedals) are excluded from the marking requirements of § 29.1555.
Par 277 413
279. § 29.674 (Amendment 29-30) INTERCONNECTED CONTROLS.
a. Explanation. A new § 29.674 is added by Amendment 29-30 which requires that the rotorcraft be capable of safe flight and landing after a malfunction, failure, or jam of any auxiliary interconnected control.
b. Procedures.
(1) Section 29.674 requires that the rotorcraft be shown to be capable of safe flight and landing after a malfunction, failure, or jam of an auxiliary control interconnected with a primary control. The section does not apply to interconnected primary controls; e.g., cyclic and collective controls.
(2) Examples of auxiliary controls covered by this section may include certain autopilot or stability augmentation or trim system components. Section 29.1309 methods may be used in determining failure effects of autopilot and stability augmentation system components.
(3) If an engine control could jam and result in a collective control jam, these controls should be designed to relieve that connection.
280. § 29.675 (Amendment 29-17) STOPS.
a. Explanation
(1) Stops are required to prevent unrestrained movements of pilot/autopilot inputs from causing interferences or overloads.
(2) The rule requires that the stop must not appreciably affect the control system range of travel due to wear, slackness, or takeup adjustments.
(3) Each stop is required to withstand loads corresponding to design conditions.
(4) In addition, each main rotor blade, if appropriate for the design, must have stops to limit its travel about its hinge points. For rotors with hingeless design, stops may be provided as appropriate to limit blade travel. Loads which may result from the blade hitting the stops (during starting or stopping the rotor, or during any large but allowable pilot control inputs such as autorotation cyclic traverse or when subjected to ground gusts, etc.) shall not overload the stops nor any rotor component.
b. Procedures.
(1) Stops are generally provided in the cockpit area and near any controllable surface end of the control system (i.e., main rotor hub, tail rotor hub, and stabilizer 414 Par 279
activators). For systems with control coupling or series actuators, stops have been located further away from the cockpit to permit increased control output during malfunction (hardover) or extreme control position cases.
(2) Location of stops in close proximity to each end of a control system will allow the stops to function most efficiently without undue deflections between the stops and the adjacent surface or the adjacent cockpit control lever or pedals. The location of stops close to the control lever or surface will help meet the requirement that the stop and its function not be appreciably affected by wear, slackness, or takeup adjustments.
Consideration should be given to limiting the total amount of takeup adjustments of both the stop and the control systems to preclude a hazardous adjustment of the control surtace range of travel.
(3) Each stop is to be substantiated for critical design conditions from either pilot effort, aerodynamic loads, hydraulic loads, or other critical loads, as applicable.
The stops can be substantiated for limit loads by the tests of § 29.681. (Deliberate misrigging of the controls on the test aircraft may be necessary to assure that the maximum limit load which the stop will be subjected to in service is applied to the stop during these tests.)
(4) The stops to limit the main rotor blade about its hinge points should be positioned to prevent the blades from striking any part of the structure, particularly during startup and shutdown operations. These stops should also limit the flapping of the static main rotor blades of the rotorcraft when they are subjected to ground gusts or rotor wash from nearby taxiing rotorcraft. Provisions should be made to prevent overloading the stops or the blade under conditions of ground gusts, rotor wash effects, or during autorotation landing flares. The need for provisions to prevent possible overloads due to ground gusts and close taxiing by adjacent rotorcraft and by autorotation landings can be determined using the instrumented flight load survey aircraft by hover-taxiing another rotorcraft near the instrumented aircraft and by conducting autorotation landing flares with the instrumented aircraft. Substantiation for the final main rotor flapping stop design can be demonstrated by similar tests.
(5) If features of design are added to the main rotor stop assembly which activate certain portions of the stop assembly only on the ground to meet the requirement that the blade not hit the droop stop during any operation other than starting and stopping the rotor, such features of design must be substantiated to reliably operate by both ground tests and flight tests, as appropriate. Wear and rigging tolerances should be considered in these demonstration tests.
Par 280 415
281. § 29.679 CONTROL SYSTEM LOCKS.
a. Explanation.
(1) Whenever a control system lock or locks are used, the standard requires design features to prevent flight or limit operation before flight begins with the lock engaged. Locks are not required by the standard.
(2) After flight begins, design features shall be used when needed to prevent possible lock engagement while the rotorcraft is in flight or ground operation.
(3) The standard applies to external control locks as well as internal locks.
b. Procedures.
(1) Locks that release or disengage automatically, as stated, may be used.
Attention should be directed to reviewing possible means of lock engagement while in flight. Fault analysis of the system should be used to ensure possible failures are determined. Design features may be used or needed to preclude this event.
(2) Manually applied and released locks may be used. Design features of the locks must prevent engagement in flight also.
(3) Any "unmistakable" warning to prevent takeoff with a lock engaged should be easily discernable during day and night operations. It should be possible to apply the lock only in such a manner that the required warning is provided. Color, location, shape (identification), and accessibility of the device or its control and legibility of any device placards or markings are important considerations in the evaluation.
(4) During a "compliance inspection," and during TIA evaluations, the locks shall be evaluated to the standards. When a lock is not automatically disengaged, the operation of the rotorcraft should be limited. Unmistakable warning may be achieved as Tollows.
(i) Prevent sufficient power for takeoff.
(її) The pilot shall be unable to move the collective control from the lowest pitch limit.
(Ill) One or more aural devices that cannot be disengaged (turned off) until all locks are removed.
416 Par 281
282. § 29.681 LIMIT LOAD STATIC TESTS.
a. Explanation.
(1) The rule requires static tests of the control system in showing compliance with limit load requirements.
(2) The tests are specified to include each fitting, pulley, and bracket of the control system being tested and to include the "most severe loading."
(3) Also, the rule requires that compliance with bearing factors (reference § 29.623) be shown by individual tests or by analyses for control system joints subject to motion.
b. Procedures.
(1) Compliance with the requirements of this rule is obtained by static tests conducted on either a static test airframe or on a prototype flying ship. In either case, conformity of the control system and related airframe is necessary to validate the tests.
(2) The rotor blades or aerodynamic surfaces may be used to react pilot effort loads through the control system or they may be replaced with fixtures. If fixtures are used, they should be evaluated for geometric and stiffness effects to assure test validity.
(3) The loads to be applied during the limit load static tests are specified in §§ 29.395, 29.397, and 29.399. The loads are applicable to collective, cyclic, yaw, and rotor blade control systems as well as any other flight control systems provided by the design.
(4) Section 29.585(e) specifies bearing factors for control system joints subject to angular motion. These factors are 3.33 for push-pull systems and 2.0 for cable systems for joints with plain bearings. For joints with ball or roller bearings, use the manufacturer's ratings.
283. § 29.683 OPERATION TESTS.
a. Explanation. The rule requires that the control system be free from jamming, excessive friction, and excessive deflection. An operational test is required in which specified loads are applied at the pilot controls and carried through an operating control system.
b. Procedures.
Par 282 417
AC 29-2B 7/30/97/
(1) Compliance with the requirements of this rule is obtained by use of a test setup similar to that used for the limit load tests of § 29.681, except the load reactions at the blades (or surfaces) must allow for movement of the blades (or surfaces) as the system is operated through its operating range.
(2) Fixtures are normally affixed to the surfaces (or replace the surfaces) to allow pulley arrangements which provide for movement under load. These fixtures should be evaluated to assure that system loads up to limit will be applied during the full range of operations of each system.
(3) Each flight control system should be operated through its entire range under a light load and under limit load. As the controls are being operated, the system should be checked for jamming, excessive friction, and excessive deflection. Excessive deflection includes deflection sufficient to contact other systems or structure. Also (in agreement with CAM 04.331/04.43.11), FAA/AUTHORITY policy has been to consider as excessive the deflection of a control system under limit load which exceeds approximately one-half of the system travel from neutral to an extreme stop. Floor panels, wall panels, and other access panels may have to be removed to permit visual checks of the entire control system. However, care should be taken when removing panels so that airframe structure is not weakened enough to deflect from its normal position when test loads are applied to the control system.
284. § 29.685 (Amendment 29-12) CONTROL SYSTEM DETAILS.
a. Explanation. The rule requires that the control system be designed to prevent chafing, jamming, and interference from cargo, passengers, loose objects, or the freezing of moisture. Specifically, means are required in the cockpit to prevent the entry of foreign objects into places where they would jam the system, and means are required to prevent the slapping of cables or tubes against other parts. Specific design considerations to prevent binding and overloads within the control system are required such as--
(1) Assure pulley-cable combinations as specified in MIL-HDBK-5 are used unless inapplicable.
(2) Assure close fitting pulley guards are provided.
(3) Assure pulley-cable alignment sufficient to prevent excessive pulley flange loads is provided.
(4) Assure fairlead-cable alignment is within 3°.
(5) Assure no clevis pins are retained only by cotter pins.
418 Par 283
(6) Assure turnbuckles do not bind other structures throughout the range of travel.
(7) Assure means for inspection of control system components are provided.
(8) Assure control system joints subject to angular motion incorporate special bearing factors, 3.33 for push-pull systems and 2.0 for cable systems.
(9) Assure that manufacturer's ratings for ball or roller bearing ratings are not exceeded.
b. Procedures.
(1) The geometry of the control system components and installations is the primary control to prevent chafing, jamming, and interference. The control system from cockpit to surface should be checked for clearances both unloaded and loaded. The control system should be checked under load during both the limit load static tests (reference § 29.681) and the operational tests of § 29.683. Location of guides or fairleads and pulleys may be used in cable systems to prevent chafing and interference with other structure. Generally, tubes should clear adjacent structure by location and design geometrical considerations. If supplemental means are provided to assure the tubes do not chafe or interfere, the means should be evaluated for possible jamming.
(2) Rubber (or other elastomeric) boots connected to both the cockpit control arm or shaft and to the floor are acceptable means to prevent the entry of foreign objects into underfloor areas where they may cause jamming of controls. Control systems should, in general, be routed around cargo compartments. If routing of the control system components is in or near cargo areas, the control system components should be protected by bulkheads, panels, or other enclosures which have sufficient strength and stiffness to prevent possible interference with the control system components when subjected to cargo loading and handling deflections.
(3) Control system details should be reviewed for possible moisture collection.
Areas should drain free. Exposed or open control areas should drain free, and areas of possible freezing moisture collection should not accumulate ice that would cause a jam of the controls. Simulated or actual ice collection on the controls may be used to prove questionable features. The areas to be considered for moisture collection include both external and internal areas where moisture may accumulate by direct impingement of water, entrapment of water particles, or condensation of moisture.
(4) The latest revisions of MIL-HDBK-5 do not explicitly give approved pulley-cable combinations, but appropriate MIL specifications are given in Chapter 8.3 for use in determining pulley-cable combinations and ratings.
(5) Provide ratings, factors, and alignment as specified.
Par 284 419