The bearing projected area is 0.312 × 2 = 0.624-square-inches. The design limit load is 3,000 pounds. The design limit bearing stress is 3,000 pounds/0.624-square-inch x
1.85 = 8,894 PSI. If a free or loose fit is not used; i.e., tighter than free fit, a bearing factor is not required. See Paragraph 251a(4) for bearing factors.
(4) Military standard part specifications, MS 21240, "Bearing, Sleeve Plain, TFE Lined," and MS 21241, "Flanged Bearing, Sleeve Plain, TFE Lined," contain allowable load ratings, static and dynamic, that apply to the particular use of the bearing. An appropriate bearing factor should be applied to the static rating. Military Specification (MIL-B-8943A, Amendment 3, "Bearing, Sleeve, Plain, and Flanged, TFE Lined" (temperature range -65° F to +250° F) shows that MS 21240 and MS 21241 sleeve bearings have been superseded by MS 1934/1 and MS 81934/2 sleeve bearings, respectively. Military Specification MIL-B-81934, Amendment 2, , "Bearings, Sleeve, Plain and Flanged, Self-Lubricating," uses TFE liners. These bearings are intended for use in a temperature range from -65° F to +325° F. Whenever a sleeve bearing is used, an appropriate bearing factor should be applied to the static rating that is contained in the specification or standard. Other sleeve bearings are contained in standards NAS 72 through NAS 77, NAS 537, and NAS 538. The installation design information is only contained in standards NAS 72 through NAS 74. These types of plain sleeve bearings are designed for clamping to the shaft or bolt with relative motion occurring on the bearing outside diameter. An appropriate bearing factor is required for the application.
(5) The minimum fitting factor 1.15, specified by § 27.625, must be applied as specified to account for load distribution at the fitting. This fitting factor need not apply to plain or journal "bearings" whose "bearing factor" exceeds 1.15.
(6) For airframe and landing gear structural joints, the manufacturer's static, non-Brinell rating of ball and roller bearings should not be exceeded. ABEC Class 1 bearings or better quality bearings may be used in airframe structural joints and landing gear; ABEC Class 3, 5, or 7 bearings should be used in rotor pivot joints. The non-Brinell rating includes consideration of the bearing factor, and no other bearing factor is necessary.
(7) A free fit was described in American Standards Association (ASA) Standard B4a-1925. The "free fit" clearances and tolerances of this old standard are now called Class RC6, Medium Running Fit, in ASA Standard B4.1, 1955. As an illustration using these standards, a 1-inch diameter shaft and a plain sleeve bearing would have a clearance ranging from 0.0014 to 0.0040 inch.
252. § 27.625 FITTING FACTORS.
a. Explanation. A 1.15 factor is specified to ensure that the calculated load and stress distribution within any fitting is conservative. Application of the factor is excluded or is an exception as stated in the rule.
320 Par 251
b. Procedures.
(1) The factor may be applied to the calculated load or stress for the fitting.
(2) The structural design substantiating data should include the fitting factor and where applicable should include, but not be limited to, the rotor system. The rotor system includes the rotor blade attachments, rotor head and hubs, and boosted control system elements. Other typical areas that may be considered are tail rotor gearbox attachment, tailboom to fuselage fittings, transmission pylon attachments, and landing gear attachment to the rotorcraft.
(3) The fitting factor is not required in the following applications:
(i) Joints such as continuous joints in metal plating, welded joints and scarf joints in wood.
(її) Elements proven by limit or ultimate load tests such as nonboosted control system parts.
(iii) Elements for which a larger load factor is used such as a casting factor, a 1.33 retention factor when required for seats and safety belts, a fatigue factor, bearing factor or special factor greater than 1.15, crash load factors that are the only design case, and crash load factors that exceed limit load factors x 1.5 x 1.15.
(iv) Elements for which the failure mode does not affect safety of flight or occupant satety.
253. § 27.629 FLUTTER.
a. Explanation. The rule requires that the rotorcraft "be free from flutter under each appropriate speed and power conditioni
b. Procedures. Freedom from flutter is to be shown for the entire rotorcraft with special attention to the blades, fins, and stabilizers.
(1) Flutter is defined as an aeroelastic instability resulting primarily from coupling of flap and pitch bending modes.
(2) Freedom from flutter may be shown by analysis or by appropriately instrumented flight flutter tests.
(3) The flight load survey proposal submitted for compliance with § 27.571 may also contain tests to fulfill compliance with § 27.629.
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АС 27-1A 7/30/97
(4) Flight loads survey data or flight flutter test data should be reviewed to ensure that excessive oscillatory deflections of rotors or surtaces will not be encountered.
(5) Sensitivity analyses should be conducted to ensure that normal wear in the pitch change mechanisms of the main rotor blades and tail rotor blades does not reduce the effective stiffnesses sufficiently to cause flutter.
322 Par 253
254.-264. RESERVED.
Intentionally Left Blank 323 (thru 333) Par 254
SECTION 16. ROTORS
265. § 27.653 (Amendment 27-2) PRESSURE VENTING AND DRAINAGE OF ROTOR BLADES.
a. Explanation. The rule requires each rotor blade to be provided with venting and drainage means (i.e., holes, etc.) or else the blade must be sealed and designed to withstand internal pressure.
b. Procedures. Although the rule provides for venting and drainage features, recently certificated blades have been designed to be sealed and to sustain the "maximum pressure differentials expected in service." For modern blade designs, the internal pressure buildup due to environmental effects and centrifugal acceleration effects (near the tip) can be readily sustained with moisture sealing accomplished. The use of sealed blades is highly advantageous and recommended because of the possibility for severe corrosion damage resulting from trapped moisture and because of the difficulty in finding internal corrosion damage by use of field level inspections.
266. § 27.659 (Amendment 27-2) MASS BALANCE.
a. Explanation. The rule requires that mass balancing of rotors and blades be provided, as necessary, to prevent excessive vibration and flutter. Further, the rule requires structural substantiation of the mass balance installation.
b. Procedures.
(1) The weight, geometry, and location of rotor and blade mass balance devices are determined as the requirements of §$ 27.571 and 27.629 are met.
(2) The structural substantiation should show static strength to meet the maneuver and gust loads of §§ 27.337, 27.339, and 27.341. In addition, the main rotor loads of § 27.547(c) should be substantiated. The fatigue strength of the mass balance devices (including structural supports) should meet the requirements of § 27.571.
(3) In addition to the appropriate strength requirements, some recent designs have included features which trap the balance weight inside a limited area even if the primary attachment means (adhesive, bolts, etc.) fail. This type of design feature is recommended because of the severe loading environment to which balance devices are subjected.
267. § 27.661 (Amendment 27-2) ROTOR BLADE CLEARANCE.
a. Explanation.
334 Par 265
(1) The rotors, main and tail, must not strike other parts of the rotorcraft during any operating condition.
(2) Section 27.411 concerns protection of the tail rotor from a ground strike.
b. Procedures.
(1) The applicant should have drawings or sketches of the rotorcraft that show an adequate minimum clearance between the rotors, main and tail, and parts of the rotorcraft. Probable flexing of the rotors should be considered in determining the minimum clearance.
(2) During parts of the FAA/AUTHORITY-conducted flight test program, frangible devices (wood dowels) or other means of measuring clearance, may be requested to confirm that the clearance shown in the drawings or sketches is adequate in certain operating conditions. Balsa wood dowels may be clamped to the aft part of the fuselage within the rotor arc. If the devices are intact after autorotation landing tests and other tests involving typical abrupt, cyclic, and rudder pedal displacements, the clearance should be satisfactory and compliance obtained. If such measuring devices are used, the type inspection report should contain a record of clearance found during the tests. It is not necessary to precisely determine the clearance but only necessary to determine "enough clearance" as stated in the rule.
268. § 27,663 (Amendment 27-2) GROUND RESONANCE PREVENTION MEANS
a. Explanation.
(1) This rule, adopted in Amendment 27-2 and revised in Amendment 27-26 requires reliability and damping action investigation for the ground resonance prevention means. The probable range of variations in service, not just the allowable range, must be established and investigated as prescribed. This probable range includes operation on the ground, and other appropriate landing surfaces applicable to the rotorcraft design shall be considered. Quantitative test data are generally obtained in compliance with this rule, but analysis or tests may be used.
(2) Appropriate maintenance information should be included in the maintenance manual (also called instructions for continued airworthiness).
(3) Paragraph 99 of this document concerns demonstrating freedom from ground resonance during certain applicant and TIA verification evaluations or tests of the rotorcraft. Section 27.241 complements the requirements of § 27.663. As noted in Paragraph 99 of this document, a specific requirement for a ground vibration survey was removed from CAR Part 6. However Section 27.663 was adopted by Amendment 27-2 to investigate possible sources of ground resonance and to assure Par 267 335
the reliability of the ground resonance prevention means, i.e., dampers, if necessary, to preclude occurrence of ground resonance. The total rotorcraft system is evaluated under this rule.
(4) Viscous dampers 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 airframe. The rule also requires investigation of the probable range of variations of these dampers, whether viscous or elastomeric, and these 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. 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 rule. An analysis report or a test proposal and subsequent test report may be used to show compliance. The probable ranges of damping restriction are an important part of the assessment. The test may be conducted in conjunction with the testing required by § 27.241. See Paragraph 99.
(i) 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.
(ili) Testing of the complete rotorcraft may be used to prove that malfunction of a single means or member of the damping system will not cause ground resonance. One method of demonstrating acceptable compliance is by removing all or most of the fluid from a damper and considering the allowable ranges of damping of the other parts of the rotorcraft damping system while operating the rotorcraft throughout the rotor speed range from start to maximum rotor speed. Investigation of elastomeric dampers may require 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 must simulate a malfunction to comply with the rule. The testing discussed, however, could be come very extensive if one were to attempt to test all combinations of all 336 Par 268
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 or no 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 must be established and investigated as prescribed and noted in Paragraph 268(b). An approved test proposal and test results report should be used for complying with § 27.663(b). If 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. See Paragraph 99, § 27.241, concerning tests and instrumentation of the test associated with complying with § 27.241. The instrumentation noted in Paragraph 99 also applies to § 27.663(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.
(5) Any maintenance procedures should be included in the "recommended" part of that manual. See Appendix A, FAR Part 27.
268A. § 27.663 (Amendment 27-26) GROUND RESONANCE PREVENTION MEANS.
a. Explanation. Amendment 27-26 clarifies that analysis as well as tests may be used to show freedom from ground resonance after malfunction or failure of a single 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 of this AC continue to apply with the addition of the need to document the establishment of the probable range of damping of ground resonance prevention means. Acceptable tire and oleo minimum Par 268 337
АС 27-1A 7/30/97 and maximum pressures as well as other identified factors should be documented in maintenance instructions if necessary to maintain the desired characteristics.
338 Par 268A
269.-277. RESERVED.
Intentionally Left Blank Par 269 339 (thru 347)
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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