(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 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 take-up 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 take-up 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.
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AC 27-1B, Chg 3 9/30/2008 AC 27.679. § 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. Three main approaches may be used to meet the requirements of this rule.
(1) The first approach is to provide a means to disengage the locks "automatically" as the pilot operates the controls. If this method is used, the means must disengage the locks in a manner so that they will not automatically re-engage during flight. The means may be physical removal of the locking device from close proximity to the control system interface, with deliberate crew action necessary to return the lock to the control system interface, or the means may be that the mechanism geometry or actions, or both, prevent the locking device from engaging in flight.
(2) The second approach is to provide locks that prevent 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.
(3) The third approach is to provide a means that warns the pilot when a flight control lock system is engaged, such as through a non-cancelable warning message or through distinct control lock markings which are clearly visible to the pilot under all conditions (day and night).
(4) Unless it can be shown that the control locks cannot be inadvertently engaged (taking into account manufacturing tolerances and maladjustments), a means should be provided that allows, in flight, a single pilot to safely disengage any flight control lock. For example, the control lock design might include frangible fittings.
Further, the RFM should describe the appropriate pilot corrective action needed to disengage the control lock system while in flight.
(5) The rotorcraft Instructions for Continued Airworthiness should include appropriate maintenance checks and procedures to be completed following modification (for example, via STC or field approval), maintenance, alignment, or adjustment that affects the flight control system locks.
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AC 27.681. § 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."
(3) Also, the rule requires that compliance with bearing factors (reference § 27.623) 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 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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AC 27.683. § 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 structure. Also, if under these limit load conditions the components deflect, the deflection would be considered excessive if there is permanent deformation of any component or supporting structure. Also any deflection that results in an uncorrected condition when the load is released, e.g., if a bellcrank is forced off-center or over-center during load and does not return to the normal position after load release is excessive deflection. Floor panels, wall panels, and other access panels may have to be removed to permit visual checks of the entire control system.
AC 27.685. § 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 Page D - 41
fairleads and pulleys may be used in cable systems to prevent chating 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.
AC 27.685A. § 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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AC 27.687.$ 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.
AC 27.691. $ 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.
AC 27.695. § 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.
(9) Section 27.1301 Function and installation.
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(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:
(1) 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 (II) A manually operated mechanical system.
(4) The power portion of the system includes:
(i) The power source (such as hydraulic pumps); and (II) 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, servo valves, servo-valve cylinder, and power actuator housings. There must be no Page D - 45
AC 27-1B, Chg 3 9/30/2008 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 exceptional piloting skill and strength in any normal maneuver for a period of time as long as that required to effect a safe landing.
The control forces should not exceed those specified in § 27.397 and flight characteristics should meet the requirements of §$ 27.141 (b) and (b)(3).
(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.
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 safe flight.
(ili) Use a single transmission-driven pump and an electrically driven pump.
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SUBPART D - DESIGN AND CONSTRUCTION
LANDING GEAR AC 27.723. § 27.723 SHOCK ABSORPTION TESTS.
a. Explanation.
(1) Limit and "reserve energy" drop tests are required as prescribed in §§ 27.725 and 27.727, respectively. These tests may be conducted on the complete rotorcraft or on units consisting of wheel, tire, and shock absorber in their proper relation. For rotorcraft with skid landing gear, the tests may be conducted on the complete rotorcraft or on a simulated fuselage with the complete skid landing gear system.
(2) The rotorcraft must be designed to limit load factors that equal or exceed the limit load factor substantiated by these drop tests. In practical application, the rotorcraft may be designed to a limit load factor, such as 2.8g. Thus, it is necessary that the limit landing load factor derived from the landing gear drop tests be equal to or less than 2.8g. If not, the rotorcraft must be redesigned for the higher load factor derived from the drop tests. It must be shown in accordance with § 27.723 that the limit load factors selected for design under § 27.473 will not be exceeded in landings with the limit descent velocity corresponding to the drop height specified in that section. In addition, reserve energy absorption capacity of the landing gear must be shown for a descent velocity of 1.22 times the limit descent velocity selected under § 27.473 by increasing the drop height to 1.5 times the "limit" drop height. The test requirements or procedures outlined in Part 27 for obtaining the landing load factors are empirical; however, these procedures are based on and supported by satisfactory experience.
(3) As stated in § 27.725(c), each landing gear unit should be tested in the attitude simulating the landing condition that is most critical from the standpoint of the energy to be absorbed by it. For wheel landing gear designs, the level landing or tail down landing and level landing with drag are generally the most critical attitude. A test of more than one attitude may be required to comply with the standard.
(4) Drop tests are required. If analytical methods and/or means are proposed by the applicant, the data presented for approval must be equal to or conservative with respect to that data obtained from physical drop tests. Section 21.21(b)(1) concerns "equivalency" determinations. Presenting an acceptable means of "equivalency" here would circumvent the necessary scrutiny of an analytical method or means and is also beyond the scope of this document.
b. Procedures. The test plan or proposal must be approved prior to official FAAJAUTHORITY tests unless satisfactory resolution of outstanding proposal or conformity inspection items can be accomplished after the test.
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(1) The following headings would be a typical table of contents for the test proposal, and a generalized explanation of the contents that may be included under each of these headings for a wheel landing gear follows.
Purpose. The regulations to which compliance is being shown by the drop tests should be identified (usually §§ 27.723, 27.725, and 27.727). Also, the rotorcraft landing gear, including the wheels and tires to be dropped, should be positively identified in the report by the manufacturer's or applicant's previously FAA/AUTHORITY-approved drawing, technical standard orders (TSO's), or other identifying FAA/AUTHORITY-approved data as applicable.
(її) Description of test setup. This section should present a description of the test fuselage or jig, method of attaching landing gear to jig, and type of accelerometer to be used to measure load factors. Proof of calibration of accelerometer should be available. The accelerometer should be mounted at the aircraft CG if a free drop of the aircraft is used or as close as practical to the centerline of the main shock absorbing component of each landing gear (oleo, strut, etc.) if each gear is tested separately. The description of the test jig, including platforms on which the gears are to be dropped, should be defined by sketches in addition to the required mathematical calculations. This data should show that the landing gear will be at the proper attitude, relative to the platform, on impact for the particular landing condition. Drawings or other approved data from which the geometry is taken should be referenced in the proposal.
The tire and oleo pressures at the time of the test should be specified. The method of measuring the deflection of the tire plus the vertical travel of the axle under impact should be described. This measurement may be accomplished by telescoping tubes attached to the point on the jig that would measure the total (tire and oleo) vertical deflection of the landing gear. Other vertical and horizontal deflections should be measured as required to determine if the landing gear has experienced permanent deformation after each drop test. The effect of surface roughness should be considered. Smooth surfaces tend to give maximum deflections where rough surfaces tend to restrict deflection and to result in maximum values of Nz. Preliminary company drop tests (at less than limit drop height) may be used to determine the critical surface roughness, or engineering evaluations may be used (without tests) when the gear configurations are such that the critical surface condition can be analytically determined (or when the load factor is shown to be negligibly affected by surface roughness).
NACA Report 1154, dated 1953, contains information that surface coefficients of friction may vary from 0.4 to 0.7. Skid landing gear standards, § 27.501(c), indicate an acceptable coefficient of friction is 0.5. A wheel landing gear design standard, § 27.479(b), indicates an acceptable coefficient of friction is 0.25. In the case of a small rotorcraft, the entire aircraft may be dropped. This may be accomplished by establishing pivot points at the main gear axles for the tail (or a point forward of the nose gear) drops and a pivot point at the tail (or nose gear) axle for the main gear drops. It is the responsibility of the applicant to distribute the aircraft inertia items, including added weight to get the proper effective drop weight (We) at the landing gear, Page D - 48