SECTION 17. CONTROL SYSTEMS
278. § 27.671 GENERAL.
a. Explanation.
(1) The rule requires basically that controls operate easily and smoothly and provide positive response of the rotorcraft from control input.
(2) In addition, the rule requires that incorrect assembly be prevented by special design features or special markings.
b. Procedures.
(1) Easy, smooth operations of controls are substantiated by the operations tests of § 27.683 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 §§ 27.141 through 27.175.
(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 §§ 27.681 and 27.683, the controls should be thoroughly reviewed for possible incorrect assembly and for any required markings supplied for compliance with this standard.
279. § 27.672 (Amendment 27-21) STABILITY AUGMENTATION.
AUTOMATIC, AND POWER-OPERATED SYSTEMS.
a. Explanation.
(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 348 Par 278
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. 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 § 27.1329(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 81.
280. § 27.673 (Amendment 27-21) PRIMARY FLIGHT CONTROL.
a. Explanation. This regulation 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 regulation was generated to clarify the application of § 27.1555 which requires markings for controls other thaniprimary flight controls or controls) whose function is obvious."
b. Procedures. The primary flight controls; i.e., cyclic stick, collective, and tail rotor pitch control pedals are excluded from the marking requirements of § 27.1555.
281. § 27.674 (Amendment 27-26) INTERCONNECTED CONTROLS.
a. Explanation. A new § 27.674 is added by Amendment 27-26 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.
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(1) Section 27.674 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 27.1309 methods may be used in determining failure effects of autopilot and stability augmentation "system" components. For components whose purposes are solely mechanical functions, the procedures associated with § 27.571 for components such as the main rotor may be used.
(3) If an engine control could jam and result in a collective control jam, the controls should be designed to relieve that connection.
282. § 27.675 (Amendment 27-16) 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 be located to 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 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 flares 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 activators). For systems with control coupling or series actuators, stops have been located farther downstream (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 stop to provide its function most efficiently without undue deflections between 350 Par 281
the stop and its adjacent surface or its 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 surface range of travel by either normal or extreme takeup adjustment.
(3) Each stop is to be substantiated for critical design conditions from either pilot effort, aerodynamic loads, hydraulic loads, and other critical loads, as applicable.
The stops can be substantiated for limit loads by the tests of § 27.681.
(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 and rotor wash from nearby taxiing rotorcraft. Provisions should be made to prevent overloading the stops or the blade under conditions of ground gusts and rotor wash effects or during autorotational landing flares. The need for provisions to prevent possible overloads due to ground gusts and close taxiing by adjacent rotorcraft and by autorotational landing can be determined using the instrumented flight load survey aircraft by hover-taxiing another rotorcraft near the instrumented aircraft and by conducting autorotational 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.
283. § 27.679 CONTROL SYSTEM LOCKS.
a. Explanation. The rule requires that if control system locks are provided, means are necessary to prevent the rotorcraft from taking off with the locks engaged or, once airborne, to prevent the locks from engaging in flight.
b. Procedures. Two main procedures may be used to meet the requirements of this rule.
(1) The first procedure is to provide a means to disengage the lock "automatically" as the pilot operates the controls. If this method is used, the means must disengage the lock in a manner that it will not automatically re-engage during flight under normal pilot operations. The means may be physical removal of the locking Par 282 351
device from close proximity to the control system interface with deliberate crew action necessary to return the device to the control system interface, or the means may be that the mechanism geometry and/or actions prevent locks from engaging in flight.
(2) The second procedure which may be used is to provide locks which so limit rotorcraft operations that it is impossible to take off with the locks engaged. Acceptable means are features which prevent engine startup or which restrict collective control operations to prevent sufficient lift for takeoff.
284. § 27.681 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."
ference $ 27.623) be shown by individual tests orby 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 efforts to ensure test validity.
(3) The loads to be applied during the limit load static tests are specified in §§ 27.395, 27.397, and 27.399. 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) Although Part 27 does not explicitly specify the bearing factors to be used in control system rotating joint tests or analyses, the factors of § 29.685 have been used in past programs. These factors are 3.33 for push-pull systems and 2.0 for cable systems for joints with plain bearings and manufacturers' ratings for ball and roller bearings.
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285. § 27.683 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.
(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 § 27.681, 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 ensure 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 structures. Also (in agreement with CAM 04.331/04.43.11), FAA/AUTHORITY policy has been to consider excessive the deflection of a control system under limit load which exceeds approximately one-half of the system travel from neutral to the extreme stop. Floor panels, wall panels, and other access panels may have to be removed to permit visual checks of the entire control system.
286. § 27.685 (Amendment 27-11) 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.
b. Procedures.
(1) The geometry of the control system components and their installations are 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 § 27.681) and the operational tests of § 27.683. Location of guides or fairleads and pulleys may be used in cable systems to prevent chafing and interference Par 285 353
AC 27-1A 7130/97 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 shalt 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 tree and areas ot 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.
286A. § 27.685 (Amendment 27-26) CONTROL SYSTEM DETAILS.
a. Explanation. Amendment 27-26 adds §§ 27.685(d), (e), and (f) for cable systems, control system joints, and bearings, which are compatible with the same pre-existing paragraphs of § 29.685 except cables of 3/32 inch diameter are allowed by this section rather than the minimum of 1/8 inch diameter required by § 29.685 for transport rotorcraft.
b. Procedures. All of the policy material pertaining to this section remains in effect. This material is supplemented with the following:
(1) The latest revisions of MIL-HDBK-5D do not explicitly give approved pulley-cable combinations, but appropriate MIL specifications are referenced in Chapter 8.3 of MIL-HDBK-5D for use in determining pulley-cable combinations and ratings.
(2) Adhere to the ratings, factors, and alignment as specified.
(3) Provide inspection means as specified for the control system.
(4) Close fitting pulley guards are required for cable systems.
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287. § 27.687 SPRING DEVICES.
a. Explanation.
(1) This standard for control systems ensures that springs and spring devices used to prevent flutter, control oscillations, or vibrations are either -- (i> Reliable (failure is extremely remote); or
(ii) The failure is not critical to the rotorcraft.
(2) Tests simulating service conditions are required in either instance.
b. Procedures.
(1) Springs and spring devices used in the control system, including balance springs, should be identified early in the certification program.
(2) Whenever a spring cannot be proven by observation or analysis that it is "not critical," then ground or flight tests may be required.
(3) Springs that are critical to safe operation may be subject to fatigue substantiation to prove they are reliable for the operating conditions imposed in service.
(4) Springs used in conjunction with hydraulic actuator spool valves may be subject to the standards of § 27.695.
288. § 27.691 AUTOROTATION CONTROL MECHANISM.
a. Explanation
(1) Rotorcraft designs generally have a main rotor blade collective pitch control system that does not have detents or other devices to limit pitch control in the control midrange. Autogyro and other rotorcraft designs may include detents or other finite position control for collective pitch control. This rule requires that the control design allows rapid entry into autorotation after a power failure.
(2) Section 27.33 contains standards concerning establishment and control of the main rotor speed limits. The standard requires flight tests and demonstrations. The standard also concerns rotorcraft design features that are related to control of the main rotor speed limits.
(3) Other design requirements for control systems are contained in § 27.685.
b. Procedures.
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(1) If high and low main rotor pitch stops are employed in the collective control and if the control may be rapidly moved from one limit to the other, compliance is shown.
(2) If detents or intermediate stops are employed, the pilot must be able to easily and readily override, disconnect, remove, or bypass the device to allow rapid autorotational entry prior to exceeding transient low speed rotor limits. An early assessment of the design may be accomplished by the flight test personnel with the evaluation completed in the Type Inspection Authorization (TIA) test program.
(3) It is acknowledged that modern rotorcraft designs may have an autorotational VNE that is lower than "power-on" VNE or normal cruise speed. For rotorcraft designs with this characteristic, the speed must be reduced after entry into autorotation. No relief from the rule is required since many phases of operation occur at speeds less than power-on VNE. For example, a critical phase of flight occurs during takeoff. Rapid entry into autorotation is essential during this phase also.
(4) The features of the autorotational control mechanism and ability to control the rotor speed within the design limits for any rotorcraft will be evaluated as an integral part of the TIA test program.
289. § 27.695 POWER BOOST AND POWER-OPERATED CONTROL SYSTEM.
a. Reference Regulations. The following sections of Part 27 are either incorporated in the provisions of § 27.695 or are otherwise applicable to power boost and power-operated control systems:
(1) Section 27.307 Proof of structure.
(2) Section 27.571 Fatigue evaluation of fight structure.
(3) Section 27.671 Control system.
(4) Section 27.681 Limit load static tests.
(5) Section 27.687 Spring devices.
(6) Section 27.685 Control system details.
(7) Section 27.861 Fire protection of structure controls and other parts.
(8) Section 27.863 Flammable fluid fire protection.
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(9) Section 27.1301 Function and installation.
(10) Section 27.1309 Equipment, systems, and installations.
b. Explanation.
(1) The rule requires an alternate system if a power boost or power-operated control system is used.
(2) The alternate system must, in the event of any single failure in the power portion of the system, or in the event of failure of all engines:
(i) Be immediately available.
(ii) Allow continued safe flight and landing.
(3) The alternate system may be:
(i) A duplicate power portion of the system; or
(il) A manually operated mechanical system.
(4) The power portion of the system includes:
The power source (such as hydraulic pumps); and
(il) Items such as valves, lines, and actuator.
(5) The failure of mechanical parts (such as piston rods and links) must be considered unless their failure is extremely improbable.
(6) The jamming of power cylinders must be considered unless their jamming is considered extremely improbable.
c. Procedures. It is assumed in the following discussion that the power boost or power-operated control system being utilized is a typical aircraft hydraulic system.
(1) The rule requires, without respect to the probability of failure, an alternate system for the power portion of the system. The power portion of the system, by example in the rule, includes hydraulic pumps, valves, lines, and actuators. It has also been interpreted to include seals, servo valves, and fittings.
(2) If a duplicate power portion of the system is used to meet the requirements of the rule, the requirements may be met by providing a dual independent hydraulic system, including the reservoirs, hydraulic pumps, regulators, connecting tubing, hoses, Par 289 357
servo valves, servo-valve cylinder, and power actuator housings. There must be no commonality in fluid-carrying components. A break in one system should not result in fluid loss in the remaining system.
(3) Dual actuators should be designed to ensure that any single failure in the duplicated portion of the system, such as a cracked housing, broken interconnecting input, or broken interconnecting output link, does not result in loss of total hydraulic system function.
(4) A manually operated mechanical system may be used as the alternate system to a single hydraulic system if, after the loss of the single hydraulic system, the pilot can control the rotorcraft without undue mental or physical fatigue in any normal maneuver for a period of time as long as that required to effect a safe landing.
(5) The substantiation of the various system components should include consideration for operation in the normal and alternate system modes.
(6) The "extremely improbable" criteria noted in § 27.695(c) for failure of mechanical parts may be satisfied by performing component fatigue testing and establishing a service life through this technique.
(7) Fatigue substantiation of the control actuator is required under § 27.571 and should consider both the stresses imposed by flight loads and the stresses imposed by hydraulic pump pressure pulses. Flight loads factored in a conservative way may be an acceptable means to take into account both effects.
(8) The possibility of jamming of the power cylinder may be shown as "extremely remote" through a failure analysis that considers every possible system component failure such as, but not limited to, ruptured lines, pump failure, regulator failure, ruptured seals, clogged filters, jammed servo valves, broken interconnecting servo valve inputs, broken interconnecting output links, etc.
(9) Three acceptable means to meet the requirements of § 27.695(a)(2) could be as follows:
(i) Provide two transmission-driven hydraulic pumps, provided the pumps are driven by the transmission during all flight conditions including autorotation.
(ii) Use two electrically-driven hydraulic pumps if electrical power is available to drive the pumps with all engines failed. If this approach is used, the battery must be capable of running both pumps plus all other required equipment necessary for continued sate flight.
(iii) Use a single transmission-driven pump and an electrically driven pump.
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290.-296. RESERVED.
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