(ii) Torque factors are also specified for reciprocating engines having two or more cylinders in § 29.361(a)(2) or § 29.361(b) of Amendment 29-26. The appropriate torque factor applies to takeoff power torque as well as maximum continuous power and other power conditions.
(2) Amendment 29-26 introduced additional turbine engine installation considerations for the following:
(i) Engine torque loads associated with emergency operation of governor-controlled turboshaft engines.
(ii) Torque reaction loads from sudden turbine engine stoppage which is applied to the engine and the engine suspension and restraint system.
(3) Paragraph AC 29.549 concerns § 29.549(c) and (e) that contains design standards for engine mounts and adjacent structure for flight and landing and also flight with 2 ½-minute OEI power rating. Amendment 29- 26 added OEI power to the standard.
(4) Section 29.547(e)(1)(ii) concerns the application of limit engine torque to design of the main rotor structure.
b. Procedures.
(1) The engine torque associated with the maximum continuous power condition should be multiplied by the appropriate torque factor to obtain the engine torque value used for structural substantiation purposes of the rotorcraft.
(2) The torque values associated with the minimum power-on RPM limit should be used. Maximum power -on speed limit will result in a lower torque value when calculating torque from design horsepower values. However, due to piston engine power output characteristics, an engine may produce a higher torque at higher engine speeds contrary to the previous statement. The torque factor should account for this characteristic.
(3) For turbine engines limit torque values are determined for the four cases cited. Two cases are related to “endurance” test standards.
(4) For sudden stoppage of turbine engines the engine manufacturer can reasonably provide engine rotating inertia and deceleration time expected in the event of sudden engine stoppage which generates these critical loads in the engine mounting and restraint system. These manufacturer’s data should be acceptable for use in complying with this part of the design standard.
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SUBPART C - STRENGTH REQUIREMENTS
CONTROL SURFACE AND SYSTEM LOADS AC 29.391. § 29.391 CONTROL SURFACE AND SYSTEM LOADS - 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 AC 29.395, AC 29.397, AC 29.399, AC 29.401, AC 29.403, AC 29.411, and AC 29.413.
AC 29.391A. § 29.391 (Amendment 29-30) CONTROL SURFACE AND SYSTEM LOADS - GENERAL.
a. Explanation. Amendment 29- 30 adds an explicit reference to § 29.427, Unsymmetrical Loads (paragraph AC 29.427), to clarify that substantiation for unsymmetrical loads is a general control surface requirement. A reference to § 29.399, Dual Control System (paragraph AC 29.399), is also added for clarification. In addition, §§ 29.401, 29.403, 29.413 were removed by this amendment since these references and requirements were adequately addressed in other standards.
b. Procedures. The referenced AC paragraphs become AC 29.395, AC 29.395A, AC 29.397, AC 29.399, AC 29.411, and AC 29.427.
AC 29.395. § 29.395 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.
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(3) Control system design feature and test requirements are found in §§ 29.671 through 29.695. Bearing factors 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/AUTHORITY approval.
(2) The rotorcraft control system may be tested to ultimate design loads or may be analyzed for the ultimate design loads. See paragraph AC 29.307.
(i) It is advisable that the applicant prepare a proposal describing the procedures and techniques to be used in the static testing of the control system which reflects compliance with the condition specified. It is further advisable that the FAA/AUTHORITY concur that the tests proposed achieve that objective. Omission of these steps may result in the need for retesting. The test results should be documented.
(ii) If tests are not conducted, a structural analysis of the control system is required. Appropriate factors from §§ 29.685(e), 29.623, and 29.625 must be used as specified. A structural analysis report should be used to document 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.
(i) The limit design loads are those loads specified in § 29.397;
(ii) 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
(iii) In lieu of a rational analysis, the limit design loads may be 0.60 of the loads specified in § 29.397.
(iv) For example, if a control surface servo tab or a small elevator is a part of the rotorcraft 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 rotorcraft. 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.
(v) If a hydraulic power actuation or boost system is part of the rotorcraft design, the design limit load for the affected parts of the control system will be the Page C - 21
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)(ii). 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 limit loads approved under § 29.395(b).
See paragraphs AC 29.681 and AC 29.683.
AC 29.395A. § 29.395 (Amendment 29-30) CONTROL SYSTEM.
a. Explanation. Amendment 29- 30 clarifies that the loads in § 29.395(b) apply to power “control” systems not just power “boost” systems; and the limit pilot forces prescribed in § 29.397 are required to be applied in conjunction with the forces from normally energized power devices. The amendment may increase required loads for systems if operational loads may be exceeded through jamming, ground gusts, control inertia, or friction. If so, the system is required to withstand 100 percent of limit pilot forces specified in § 29.397, rather than 60 percent of the limit pilot forces as specified previously.
b. Procedures. The procedures of paragraph AC 29.395 continue to apply except that the increased loads in new paragraph § 29.355(b)(4) of 100 percent of limit pilot forces are specified for systems where operational loads may be exceeded by jamming, ground gusts, control inertia, or friction.
AC 29.397. § 29.397 (Amendment 29-12) 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) of the rule.
b. 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.
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AC 29.399. § 29.399 DUAL CONTROL SYSTEM.
a. 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) 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.
AC 29.401. § 29.401 (Amendment 29-4) 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 structural 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.
b. Procedures.
(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.
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(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).
(i) 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.
(ii) The rotor operating speed limits shown during the official FAA/AUTHORITY flight tests must include the noted 5 percent margin with respect to the design speeds.
(iii) 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 bearings 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 is a means of proving that use of a particular bearing is satisfactory.
(v) The analysis must include appropriate special factors, casting factors, bearing factors, and fitting factors prescribed by §§ 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.
AC 29.401A. § 29.401 (Amendment 29-31) AUXILIARY ROTOR ASSEMBLIES.
a. Explanation. Amendment 29- 31 removed this section since the requirements are adequately addressed in §§ 29.337, 29.339, and 29.341.
b. Procedures. The policy material pertaining to this section is retained as supplemental information.
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AC 29.403. § 29.403 AUXILIARY ROTOR ATTACHMENT STRUCTURE.
a. Explanation.
(1) The auxiliary rotor attachment structure(s), which is considered to include gear boxes, must be designed to withstand design limit loads that occur in flight and on landing. These design loads that generally consist of the following must be established for the particular flight and landing condition under consideration.
(i) Inertia loads generated by linear and angular accelerations of the auxiliary rotors and their gear boxes, combined with
(ii) Thrust and torque loads developed by the auxiliary rotors.
The linear and angular acceleration loads imposed by the weight of the tail rotor and gearbox are generally derived from airframe loads data. Thrust and torque output of the tail rotor are derived during external aerodynamic and landing loads development for pertinent flight and landing conditions.
(2) General rules related to proof of structure loads and factor of safety are §§ 29.307, 29.301, 29.303, and 29.305.
b. Procedures.
(1) The angular and linear acceleration loads combined with appropriate tail rotor thrust and torque for the critical conditions shall be imposed on the tail rotor gearbox mount lugs, the airframe mounting structure, and the attaching hardware.
(2) The yaw and maximum power climb conditions are generally critical.
Landing and maneuvering conditions with and without power may also impose high inertia and rotor thrust and torque loads on the attachment structure.
(3) The derivation of the loads and conditions are too extensive to include here.
Additional information can be found in the U.S. Army Material Command Report AMCP 706-201, “Engineering Design Handbook: Helicopter Engineering, Part One, Preliminary Design.”
AC 29.403A. § 29.403 (Amendment 29-31) AUXILIARY ROTOR ATTACHMENT STRUCTURE.
a. Explanation. Amendment 29- 31 removed this section since the requirements are adequately addressed in §§ 29.337, 29.339, and 29.341.
b. Procedures. The policy material pertaining to this section is retained as supplemental information.
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AC 29.411. § 29.411 GROUND CLEARANCE: TAIL ROTOR GUARD.
a. Explanation.
(1) The rule requires specific protection to prevent the tail rotor from contacting the landing surface during a normal landing if it is possible that the tail rotor will contact the surface. The rule states that it must be impossible for the tail rotor to contact the surface during a normal landing.
(2) If a guard is required, the guard and its supporting structure must withstand suitable design loads.
(3) Section 29.501(c)(1) contains skid landing gear drag requirements that may be applied to the guard design loads.
b. Procedures.
(1) The applicant may submit sketches or drawings showing probable clearance with typical level landing surfaces during normal landings. Typical attitudes such as nose high autorotation, or autorotation with power-on landing, or other possible tail low attitudes should be investigated. If the drawings or sketches reveal that it is not likely the tail rotor will contact the landing surface, this minimum clearance with the landing surface may be confirmed during official FAA/AUTHORITY flight tests, such as HV and landing tests. The clearance may be confirmed by having a frangible device of suitable length (i.e., a balsa wood dowel) extending beyond the guard and attached to the tail rotor guard or other appropriate fuselage part. If the device is not damaged, broken, or no contact is made with the surface, compliance has been demonstrated.
(2) If it is possible for the tail rotor guard to contact the landing surface suitable design loads must be established for the guard. ANC -2a dated March 1948, “ANC Bulletin Ground Loads,” paragraph 6.4, entitled “Tail Bumper Criteria,” is an acceptable means of deriving the rotorcraft kinetic energy that shall be absorbed by the guard. This method is noted here for convenience.
(i) The tail rotor guard shall be able to absorb the kinetic energy of the rotorcraft in its most unfavorable CG position in the tail down landing attitude. The kinetic energy that the tail rotor guard shall be capable of absorbing must be determined as follows:
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WVS2 KY2 KE = ________ X ____________ 2g (KY2 + 1B2) where-- VS = vertical speed ft/sec, derived from § 29.725(a) KY = pitching radius of gyration - ft. from pitching axis 1B = distance from most critical CG location to the guard or bumper contact point - ft.
W = gross weight less rotor lift from § 29.473(a) - lbs.
G = 32.2 ft./sec2
(ii) 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.
(iii) 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 the guard and the rotorcraft tail (see § 29.305).
(3) Substantiation of 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 rotorcraft because of its design features.
AC 29.413. § 29.413 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 application 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 CG range and increasing the airspeed of modern designs.
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(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.
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.
(i) Limit loads of 15 pounds per square foot will apply up to approximately 90-knot design airspeed. Above a 90- knot design airspeed (VD), the coefficient (CN = 0.55) imposes higher limit loads on the surface.
(ii) In addition, combined maneuvering and gust loads may impose the highest limit loads on the control surfaces of rotorcraft. 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.
(iii) 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.
(i) The spanwise load may be rectangular or other acceptable conservative distributions may be used. The method developed by O. Schrenk in NACA TM 948, 1940, is 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.
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(ii) The chordwise distribution appropriate for the aerodynamic shape should be used.
(iii) 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 also.
(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 prescribed. If analysis is used, fitting factors and other appropriate factors prescribed by the rules of §§ 29.625, 29.621, and 29.623 will be required in the analysis.
AC 29.413A. § 29.413 (Amendment 29-31) STABILIZING AND CONTROL SURFACES.
a. Explanation. Amendment 29- 31 removed this section since the requirements are adequately addressed in §§ 29.337, 29.339, and 29.341.
b. Procedures. The policy material pertaining to this section is retained as supplemental information especially as reference material for paragraph AC 29.341 (§ 29.341).
AC 29.427. § 29.427 (Amendment 29-31) UNSYMMETRICAL LOADS.
a. Explanation. Amendment 29- 30 added the standard and Amendment 29-31 amended it. Minimum unsymmetrical design loads are specified for horizontal tail surfaces and also vertical tail surfaces whenever they support the horizontal tail surfaces.
(1) Loads are derived by rational analysis, or for earlier certification bases, the prescribed empirical loads of § 29.413 may be used. Section 29.413 was removed by Amendment 29-31 since the requirements are adequately addressed in §§ 29.337,
29.339, and 29.341.
(2) Rational loads, appropriate for the aerodynamic surfaces, should be distributed according to the standard.
(3) When vertical tail surfaces support the horizontal tail surfaces, the vertical tail surfaces and supporting surfaces are required to support the critical combination of vertical and horizontal surface loads distributed as shown.
b. Procedures. Two basic loading conditions are required by § 29.427 for each of the two basic empennage configurations shown.
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(1) Horizontal surfaces supported by the tail boom or fuselage. Structural substantiation should be provided for all six combinations shown in figure AC 29.427-1.
All of these empirical loading distributions should be used unless rational analysis shows one or more of each set of conditions to be non-critical or equal or more realistic distributions are substantiated. Rectangular spanwise air load distribution should be used unless more rational distribution is substantiated. If end plates are used, the air loads should be distributed accordingly.
(i) First unsymmetrical loading condition:
(A) 100 percent of the flight load is applied to one side of the plane of symmetry; and 0 percent of the flight load is applied on the other side of the plane of symmetry.
(B) For surfaces with end plates or other similar devices, the load distribution will be changed accordingly.
(ii) Second unsymmetrical loading condition:
50 percent of the flight load on one side of the plane of symmetry acting up; and 50 percent of the flight load on the other side of the plane of symmetry acting down.
(2) Horizontal surfaces supported by a vertical surface. Structural substantiation should be provided for all six combinations shown in figure AC 29.427-2.
All of these empirical loading distributions should be used unless rational analysis shows one or more of each set of conditions to be non-critical or equal or more realistic distributions are substantiated. Rectangular spanwise air load distribution should be used unless more rational distribution is substantiated. If end plates are used, the air loads should be distributed accordingly.
(i) First unsymmetrical loading condition:
100 percent of the flight load on one side of the plane of symmetry; and 0 percent of the flight load on the other side of the plane of symmetry.
(ii) Second unsymmetrical loading condition:
50 percent of the flight load on one side of the plane of symmetry acting up; and 50 percent of the flight load on the other side of the plane of symmetry acting down.
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