that the belt has been renovated and show the date of renovation. The use of letter "R" followed by the date would be acceptable. This marking could be in the form of an indelible ink stamping or cloth label stitched to the web-
362. FLARES. Parachute flares are made of materials which are subject to decomposition upon aging. Humidity affects the small igniting charge and also the materials of the candle (illuminant). Hence, the percentage of misfires in old flares is likely to be quite high. To assure unfailing performance of flares, periodically inspect the flare installation. Inspect the entire system, starting at the release mechanism in the cockpit and ending at the flare. Only a qualified person should attempt such inspection, since inadvertent discharge of such pyrotechnics may cause serious damage. Past experience has indicated that it is advisable to return all electrically or pistol-operated flares to the manufacturer for reconditioning within a maximum period of 3 years, and that for mechanically operated flares, a maximum period of 4 years is recommended.
363. OXYGEN SYSTEMS.
a. General. The following instructions are to serve as a guide for the inspection and maintenance of aircraft oxygen systems. The information is applicable to both portable and permanently installed equipment.
(1) Aircraft Gaseous Oxygen Systems. The oxygen in gaseous systems is supplied from one or more high- or low-pressure oxygen cylinders. Since the oxygen is compressed within the cylinder, the amount of pressure indicated on the system gauge bears a direct relationship to the amount of oxygen contained in the cylinder. The pressure-indicating line connection is normally located between the cylinder and a pressure-reducing valve.
NoTE: Some of the gaseous oxygen systems do not use pressure-reducing valves. The high pressure is reduced to a useable pressure by a regulator. This regulator is located between the high- and low-pressure system.
(2) Aircraft Liquid Oxygen Systems. Thus far it has not been a practice to use liquid oxygen in Par 361 AC 43.13-1A 1972 civil aircraft due to its complexity. This however, may change at any time as technological advances are made.
(3) Portable Oxygen Systems. The three basic types of portable oxygen systems are: demand, pressure demand, and continuous flow. The component parts of these systems are identical to those of a permanent installation with the exception that some parts are miniaturized as necessary. This is done in order that they may be contained in a case or strapped around a person's shoulder. It is for this portability reason that it is essential special attention be given to assuring that any storage or security provision for portable oxygen equipment in the aircraft is adequate, in good condition, and accessible to the user.
NOTE: Check portable equipment including its security provisions frequently, as it is more susceptible to personnel abuse than a permanently installed system.
b. Inspection. Hands, clothing, and tools must be free of oil, grease, and dirt when working with oxygen equipment. Traces of these organic materials near compressed oxygen may result in spontaneous combustion, explosions, and/or fire.
(1) Oxygen Tanks and Cylinders. Inspect the entire exterior surface of the cylinder for indication of abuse, dents, bulges, and strap chafing.
(a) Examine the neck of cylinder for cracks, distortion, or damaged threads.
(b) Check the cylinder to determine if the markings are legible.
(c) Check date of last hydrostatic test. If the periodic retest date is past, do not return the cylinder to service until the test has been NOTE: This test period is established by the Department of Transportation in the Code of Federal Regulations, Title 49, Chapter I, Paragraph 173.34.
(d) Inspect the cylinder mounting bracket, bracket hold-down bolts and cylinder holding straps for cracks and deformation, cleanliness, and security of attachment.
(e) In the immediate area where the cylinder is stored or secured, check for evi- Chap 8
SECTION 9. CONTROL SURFACE AND SYSTEM LOADS
152. 1 29.391 (through Amendment 29-19) GENERAL.
a. Explanation. This general rule concerns requirements for design loads of tail rotors, control or stabilizing surfaces, and their control system.
b. Procedures. The design criteria and/or the design loads report must contain the loads dictated by the referenced rules. See paragraphs 153, 154, 155, 156, 157, 158, and 159 of this document.
153. § 29.395 (through Amendment 29-19) CONTROL SYSTEM.
a. Explanation. Control system design loads and the application of these loads are contained in this rule.
(1) Paragraph (a) of the rule specifies the way or means of reacting the design loads specified in §$ 29.397 and 29.399 (for dual control systems).
The design loads must be imposed on any locks and stops and irreversible mechanisms in the control system. Both rotor blade horns and control surface horns must react, without failure, the specified loads while the controls are in critical positions.
(2) Paragraph (b) of the rule specifies application of limit pilot forces or of the maximum loads that can be obtained in normal operation, including any single power boost system failure, whichever is greater. However, minimum limit pilot force 0.60 of the loads specified in §§ 29.397 and 29.399, may be used, as specified, in parts of the primary control system that are not stiff enough to react to the loads specified in the first part of paragraph (b) of the rule. Note the objective for a rugged control system.
(3) Control system design feature and test requirements are found in §§ 29.671 through 29.695. Bearing faotors and fitting factors are specified in §$ 29.623 and 29.625, respectively.
b. Procedures.
(1) The design criteria and/or a design loads report that includes the primary control system design loads should be submitted for FAA approval.
(2) The rotorcraft control system may be tested to ultimate design loads or may be analyzed for the ultimate design loads. See paragraph 124 of this document.
(1) It is advisable that the applicant prepare a proposal control system which reflects compliance with the condition specified. It is describing the procedures and techniques to be used in the static testing of the Pa conour that the teste further advisable that the FAA conour that the tests proposed achieve that objective. Omission of these steps may result in the need for retesting. The test results should be documented.
Chap 2 Par 152 249
AC 29-2, CHG 1 11/19/84 * (11) If tests are not conducted, a structural analysis of the control system is required. Appropriate factors from §§ 29.685 (e), 29.623, and
39.025 must be used as specified. A structural analysis report should be used to jocument compliance with § 29.685(d)(1) and (4), and § 29.685(f).
(3) If a part of the control system is not stiff or rigid enough to react the design loads specified in § 29.397, that part of the system may be substantiated for lower loads as prescribed.
(1) The limit design loads are those loads specified in § 29.397;
(11) The limit design loads are the maximum that can be obtained in normal operation, including any single power boost system failure, except for objectives stated for a rugged system; and
(111) In lieu of a rational analysis, the limit design loads may be
0.60 of the loads specified in $ 29.397.
is a part of the helicopter design, the control system for this part must be stiff enough to react the control surface loads without failure and to provide enough surface deflection to control the helicopter. These limit loads may be 60 pounds fore and aft and 40 pounds laterally on the cyclic control stick in lieu of a rational analysis and may be the maximum loads that can be obtained in normal operation.
If a hydraulic power actuation or boost system is part of the helicopter design, the design limit load for the affected parts of the control system will be the maximum output force of the boost at normal operating pressure added to the limit design loads resulting from the loads specified in § 29.397.
If a single failure in the power portion of the hydraulic system results in actuator forces that exceed the maximum output force at normal operating pressure, the highest output loads must be used as noted in subparagraph
(3)(11). This hydraulic system failure standard is specified in § 29.695(a)(1) as well.
(4) Controls proof and operation test is required by $$ 29.307(b),
29.681, and 29.683. This test is conducted using the design 1imit loads approved, under § 29.395 (b). See paragraphs 282 and 283 of this document.
154. § 29.397 (through Amendment 29-19) LIMIT PILOT FORCES AND TORQUES.
a. Explanation. Design forces are contained in the rule.
(1) Primary controls, pilot and copilot, must be designed for the limit pilot forces specified in paragraph (a) of the rule.
(2) For other operating controls, such as flap, tab, stabilizer, rotor brake, and landing gear, design limit forces are specified in paragraph (b).
Chap 2 250 Par 153
D. Procedures.
(1) Design loads specified in the rule must be used in required structural tests and in any structural strength analysis of the control systems submitted in compliance with other rules.
(2) Operation tests of the control systems noted in other rules require application of these forces also.
155. § 29.399 (through Amendment 29-19) DUAL CONTROL SYSTEM.
Explanation. Design limit loads are specified for dual control systems. Pilot effort forces applied in opposition and in the same direction are required for dual control systems.
b. Procedures. 1
(1) Design loads specified in the rule must be used in required structural tests and in any structural strength analysis submitted for compliance with the other rules.
(2) Operation tests of the control systems, noted in other rules require application of these forces also.
156. § 29.401 (through Amendment 29-19) AUXILIARY ROTOR ASSEMBLIES.
a • Explanation.
(1) For rotorcraft equipped with auxiliary rotors, normally called tail rotors, an endurance test is required by § 29.923 and struotural strength substantiation is required. Section 29.401(b) specifically refers to structural strength substantiation for centrifugal loads resulting from maximum design rotor rpm. Due to the pitch feathering requirements, auxiliary rotors typically have detachable blades.
(2) The rotor blade structure must have sufficient strength to withstand not only aerodynamic loads generated on the blade surface, but also inertial loads arising from centrifugal, coriolis, gyroscopic, and vibratory effects produced by this blade movement. Sufficient stiffness and rigidity must be designed into the blades to prevent excessive deformation and to assure that the blades will maintain the desired aerodynamic characteristics. As a design objective the structural strength requirements should be met with the minimum material. Excess blade weight imposes extra centrifugal loads that may increase the operating stress levels. Blade weight and strength should be optimized.
Even though a structural strength analysis for the blade design loads is required, a flight load survey and fatigue analysis are also required by § 29.571.
(3) Section 29.1509 defines the design rotor speed as that providing a 5 percent margin beyond the rotor operating speed limits.
Chap 2 Par 154 251
AC 29-2, CHG 1 11/19/84
b. Procedure. (1) The endurance, and control parol (1) The endurance tests prescribed by $$ 29.923 and 29.927 require achieving certain speeds, power, and control displacement for the auxiliary (tail) rotor as well as the main rotor. The parts must be serviceable at the conclusion of the tests.
(2) Structural substantiation of the auxiliary (tail) rotor is required to assure integrity for the minimum and maximum design rotor speeds and the maximum design rotor thrust in the positive and negative direction. Thrust capability of the rotor should offset the main rotor torque at maximum power as required by § 29.927 (b).
(1) The maximum and minimum operating rotor speed, power-off, is 95 percent of the maximum design speed and is 105 percent of the minimum design speed, respectively.
stant test auth we on red pros (11) The rotor operating speed limits shown during the official FAA flight tests must include the noted 5 percent margin with respect to the design
(111) The auxiliary rotor generally has a positive and negative pitch limit that assures adequate directional control throughout the operating range of the rotorcraft. The power-off rotor speed limits are generally broader than the power-on rotor speed limits because of the required autorotational rotor speed characteristics. Thus, the auxiliary rotor design conditions concern the maximum and minimum design rotor speeds in conjunction with the maximum positive or negative pitch thrust as appropriate. Thrust capability and precone angle of the rotor, if any, will significantly influence the rotor design loads. The variations in rotor design features and an example of substantiation would be too lengthy to include here. However, ANC-9, "Aircraft Propeller Handbook," contains principles that may be applied to tail rotor designs. Tail rotors may be considered a special propeller design.
(iv) Bearings are generally used in the tail rotor installation to allow flapping and feathering motion of the blades. The bearing manufacturer's ratings of these bearings must not be exceeded. Bearings generally used in main and tail rotors are classified as ABEC Class 3, 5, or 7. Class 7 is the highest quality presently available. Satisfactory completion of the endurance tests of §§ 29.923 and 29.927 18 a means of proving that use of a particular bearing 18 satisfactory.
(v) The analysis must include appropriate special factors, casting factors, bearing factors, and fitting factors prescribed by 99 29.619, 29.621,
29.623, and 29.625, respectively. The fitting factor of 1.15 must be applied in the analysis of the tail rotor installation.
Chap 2 252 (thru 254) Par 156
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.
§ 29.241. See paragraph 99.
(1) Analysis and tests may be used.
(11) 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.
(111) 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 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 helicopters 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. * Chap 2 415 Par 268
AC 29-2A, CHG 2 9/24/91
(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.
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.
269.-275. RESERVED.
416 (thru 432) Chap 2 Par 268
Advisory U.S. Depariment of transportation Circular Federal Aviation Administration Subject: CERTIFICATION OF TRANSPORT Date: 9/24/91 AC No: 29 - 2A CATEGORY ROTORCRAFT Initiated by: ASW-110 Change: 2
1. PURPOSE: This change revises existing material in 17 paragraphs and adds new material for 41 paragraphs previously shown as "RESERVED."
The change number and date of the changed material are carried at the top of each page. The asterisks (*) in the right and left margins indicate the beginning and end of each change. Rearranged pages having no new material also carry the change number and new date. Pages having no changes retain the same heading information. In paragraphs that are entirely new, asterisks (*) appear only in the margins at the beginning and the end of each new page. In addition, several paragraphs were renumbered for better continuity between AC's 27-1 and 29-2A.
2. PRINCIPAL CHANGES.
a. Paragraphs 72, 80, 141, 206, 245, 268, 304, 373, 498, 621, 637, 656, 726, 748, 775, 776, and 786 are revised.
b. New paragraphs 87, 302, 303, 305, 316, 317, 318, 319, 330, 343, 347, 360, 449, 462, 549, 550, 562, 567, 568, 569, 570, 584, 585, 586, 587, 588, 589, 591, 592, 593, 594, 595, 596, 703, 705, 707, 708, 718, 725, and 785 are added to Chapter 2.
c. New paragraph 788 is added to Chapter 3.
d. Paragraphs listed below have had the paragraph numbers changed.
Existing Numbers New Numbers 707 (Proposed) 708 708 (Proposed) 707
AC 29-2A, CHG 2 9/24/91
PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated III 4/24/89 III thru XIV 9/24/91 IV 9/16/87 V thru XII 4/24/89 XIII 9/16/89 XIV 4/24/89 135 thru 138 9/16/87 135 thru 137 9/24/91 | 138 9/16/87 149 thru 152 9/16/87 |149 9/16/87 150 thru 152 9/24/91 163 and 164 (thru 174) 9/16/87 163 9/16/87 164 (thru 174) 9/24/91 237 thru 250 9/16/87 237 9/16/87 238 thru 240 (thru 248) 9/24/91 249 9/24/91 250 9/16/87 317 and 318 4/24/89 317 9/24/91 318 4/24/89 387 and 388 (thru 398) 4/24/89 387 9/24/91 388 (thru 398) 4/24/89 411 thru 432 9/16/87 411 9/16/87 412 thru 416 (thru 432) 9/24/91 491 4/24/89 491 4/28/89 492 (thru 498) 9/16/87 492 thru 504 (thru 530) 9/24/91 499 thru 500 (thru 530) 9/16/87 537 and 538 (thru 540) 4/24/89 537 thru 540 9/24/91 573 and 574 (thru 578) 9/16/87 573 9/24/91 574 (thru 578) 9/16/87 581 and (thru 596) 9/16/87 581 thru 584 (thru 596) 9/24/91 599 and 600 (thru 602) 9/16/87 599 9/16/87 1600 thru 602-2 9/24/91 607 9/16/87 607 thru 620 (thru 634) 9/24/91 608 (thru 634) 4/24/89 707 9/16/87 707 9/16/87 708 4/24/89 708 and 708-1 (and 708-2) 9/24/91 751 and 752 (thru 770) 4/24/89 1751 4/24/89 752 thru 754 (thru 770) 9/24/91 789 and 790 9/16/87 789 9/24/91 790 9/16/87 881 and 882 (thru 888) 4/24/89 1881 4/24/89 882 thru 886 (thru 888) 9/24/91 889 and 890 (thru 900) 4/24/89 4/24/89 890 (thru 900) 9/24/91 903 and 904 (thru 940) 4/24/89 903 4/24/89 904 thru 906 (thru 920) 9/24/91 921 thru 930 (thru 940) 9/24/91 2
where, Vs = vertical speed ft/seo, derived from § 29.725(a) Ky = pitching radius of gyration - ft. from pitching axis 1b = distance from most critical c.g. location to the guard or bumper contact point - ft.
# = gross weight less rotor lift from § 29.473(a) - lbs.
• = 32.2 ft./gec.2
(11) Other, more recent, analytical techniques (most utilizing computer programs) may, of course, be used rather than the ANC-2a means after proper substantiation for applicability and validity.
The tail rotor guard shall not fail when the limit and ultimate load, which is derived from a combination of the limit kinetic energy and the guard resulting limit deflection required to dissipate the energy, is imposed on
(3) substantiation or the guard, skid, or bumper for the design loads derived may be accomplished by test or analysis as stated in § 29.307(a).
(4) Several rotorcraft tail rotor guards are installed solely for the protection of ground personnel from the rotating tail rotor. For guards installed for this purpose, the applicant should use prudent and reasonable design loads and features. Such guards should not present a hazard to the helicopter because of its design features.
* 159. § 29.413 (through Amendment 29-19) STABILIZING AND CONTROL SURFACES.
a. Explanation. Minimum design loads are specified for stabilizing as well as control surfaces.
(1) Paragraph (a) of the rule requires application of minimum empirical design loads, application of critical maneuvering loads, and applioation of critical maneuvering loads combined with vertical or horizontal gust loads (30 feet per second per $ 29.341).
(2) Paragraph (b) requires load distributions that closely simulate actual pressure distributions. Both spanwise and chordwise distributions are intended.
(3) These surfaces are used for stability and control thereby hopefully extending the c.g. range and increasing the airspeed of modern designs.
(4) To "closely simulate actual pressure condition" on the surfaces, unsymmetrical loads are also required on horizontal surfaces. An arbitrary distribution, if conservative, may be used.
(5) It is noted § 29.571 requires fatigue substantiation of the flight structure which will include control and stabilizing surfaces.
(6) If the surface is controllable, a proof and operation test of-the surface control system is required by §$ 29.681 and 29.683.
Chap 2 Par 158 257
AC 29-2, CHO 1 11/19/84 * b. Procedures. Modern rotorcraft designs have generally employed a fixed or a wholly movable, not split or divided, stabilizing or control surface.
(1) Design Loads.
(1) Limit loads of 15 pounds per square foot will apply up to approximately 90-knot design airspeed. Above 90 knots design airspeed (Vd), the coefficient (CN = 0.55) imposes higher limit loads on the surface.
highest limit loads on the control surfaces of helicopters. This is attributed to the increase in speed (horizontal gust) and to the change in angle of attack and change in airspeed (vertical gust). Imposing the horizontal gust (30 feet per second or 17.8 knots) on the surface in combination with 130-knot design speed results in a 30 percent increase in the design load. The gust conditions cause a significant increase in design loads due to a change in angle of attack, with a change in resultant airspeed, or due to the increase in airspeed.
(111) The applicant may choose to derive the limit loads using maximum aerodynamic coefficients for the surface under consideration at the maximum design airspeed combined with a 17.8-knot gust. This would be acceptable provided these design loads exceed the minimum loads derived from a CN = 0.55 at design airspeed or exceed 15 pounds per square foot load on the surface.
(2) The load distribution on the surface should closely simulate actual pressure distributions.
(1) The spanwise load may be rectangular or other acceptable conservative distributions may be used. The method developed by 0. Schrenk in NACA IM 948, 1940, 1s an acceptable method for approximation of spanwise distribution.
NOTE: The method is valid for aspect ratios of 5 to 12 and for rectangular planforms such as used on rotorcraft, other planforms may be acceptable as prescribed in the TM.
(11) The chordwise distribution appropriate for the aerodynamic shape should be used.
(111) The flight load survey conducted under § 29.571 may be used to confirm design parameters and possible load distribution data. On controllable surfaces, the pitching moment (control loads) is measured for fatigue substantiation of the control system. The control stabilizing surfaces are subject to loads measurement and possible fatigue tests for fatigue substantiation
(3) Proof of the structure for the required loads is specified in §§ 29.301, 29.303, 29.305, and 29.307. Tests or analysis may be used as presoribed by the rules of $$ 29,625, 29,621, and 29.623 will be required in the prescribed. If analysis is used, fitting factors and other appropriate factors analysis.
160.-169. RESERVED.
Chap 2 258 (thru 272) Par 159
SECTION 10. GROUND LOADS * 170. $ 29.471 (through Amendment 29-19) GENERAL.
a. Explanation. This regulation specifies that limit ground loads must be considered which are:
(1) External loads caused by landing (ground) conditions and by ground taxiing loads as specified in § 29.235.
(2) Loads considering the rotorcraft structure as a rigid body.
(3) Loads in equilibrium with linear and angular inertia loads.
(4) The critical center of gravity 'must be selected so that the maximum design loads are obtained in each landing gear element."
b. Procedures.
(1) The standards to be considered are specified in $$ 29.473 through
29.511. These associated standards cover landing gear arrangements, landing conditions, and ground handling conditions.
(2) Drop tests are required for determination of landing load factors.
See paragraph 298 of this document.
(3) The application of the design loads derived from the landing load factors will be as specified for each element affected by landing or ground handling loads.
(4) During the applicant's flight test program, the ground, landing, and taxiing load factors may be monitored to assure the design load factors used are adequate. See paragraph 97 of this document for § 29.235 policy.
171. § 29.473 (through Amendment 29-19) GROUND LOADING CONDITIONS AND ASSUMPTIONS.
a. Explanation. The rotorcraft is to be designed for the maximum weight.
A rotor lift of two-thirds of the design maximum weight may be used. The minimum limit landing load factor is determined by the drop tests of § 29.725.
Provisions are made for supplementary energy absorption devices that have triggering mechanisms.
b. Procedures. Loads for the landing conditions are derived considering mass (equal to the maximum weight) and rotor lift (equal to two-thirds of the maximum weight) acting through the center of gravity throughout the landing impact. Unbalanced external loads resulting from asymmetric loading conditions are reacted as specified in the individual subparagraphs. • Chap 2 Par 170 213