90° 90° YAW DEGREES 10° MAX.
0° AIRSPEED 30 KNOT LESSER OF
SIDEWARD VH OR VNE FLIGHT FIGURE 140-1 SAMPLE YAW/FORWARD SPEED DIAGRAM 246 Par 140
(vili) During flight test evaluations, yaw angles have been measured using a yaw angle probe (swiveling vane type) on a nose boom. Both a visual readout for the pilot and a record, such as an oscillograph trace, have been used. This test may be conducted in the flight test program or in the flight load survey program. This record should confirm the yaw angle used in design as conservative with respect to operational and actual flight characteristics. This test is not a requirement however.
140A. § 29.351 (Amendment 29-30) YAWING CONDITIONS.
a. Explanation. Amendment 29-30 adds maximum sideslip angles to the existing § 29.351 for structural design purposes. The standard should apply to power-on conditions; not power-off, since Vu is a part of the standard. For airspeeds up to 0.6 VNE, sideslip angles larger than 90° (or sideward flight) need not be considered. For airspeeds at VNE or Vu (whichever is less), sideslip angles larger than 15° need not be considered.
b. Procedures.
(1) All of the policy material pertaining to this section remains in effect with the addition of the maximum sideslip limits of 90° and 15° specified above. The rotorcraft does not need to be capable of attaining these conditions. A revised yaw/forward speed diagram is presented in Figure 140-2.
(2) FAR § 29.351(b)(1) incorrectly references § 29.395(a) for maximum pilot forces. The correct reference should be § 29.397(a).
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90°
YAW DEGREES (LONGITUDINAL AXIS) • 15° MAX.
0.6 VNE AIRSPEED
LESSER OF
VA OR VNE FIGURE 140-2. SAMPLE YAW/FORWARD SPEED DIAGRAM 248 Par 140
140B. § 29.351 (Amendment 29-40) YAWING CONDITIONS.
a. Explanation. Amendment 29-31 deleted FAR § 29.403 and § 29.413 since the references and requirements are adequately addressed in §§ 29.337, 29.339, and
29.341. Paragraphs 157 and 159 of this AC are retained as guidance information.
b. Procedures. All of the policy material pertaining to this section remains in effect except for references to §§ 29.403 and 29.413 in Paragraphs 140a(3) and (4) of this AC.
141. § 29.361 (Amendment 29-26) ENGINE TORQUE.
a. Explanation.
(1) The rotorcraft should be designed for limit engine torque values, as prescribed by the rule, to account for maximum engine torque, including certain transients and torsional oscillations. The rule recognized that reciprocating (piston) engines generate higher torque oscillations than turbine engines.
A factor of 1.25 applies to maximum continuous power for turbine engines. Section 29.923 refers to torque output and § 29.927 (b) refers to other torque output conditions for use in an "endurance test."
(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 206 of this document 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) (il) concerns the application of limit engine torque to design of the main rotor structure.
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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.
142.-151. RESERVED.
250 (thru 260) Par 141
SECTION 9. CONTROL SURFACE AND SYSTEM LOADS
152. § 29.391 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.
152A. § 29.391 (Amendment 29-30) GENERAL.
a. Explanation. Amendment 29-30 adds an explicit reference to § 29.427, Unsymmetrical Loads (AC Paragraph 162), to clarify that substantiation for unsymmetrical loads is a general control surface requirement. A reference to § 29.399, Dual Control System (AC Paragraph 155), 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 153, 153A, 154, 155, 158, and 162.
153. § 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.
(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.
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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 124 of this document.
(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.
(її) 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;
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.
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 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 262 Par 153
maximum output force at normal operating pressure, the highest output loads must be used as noted in Subparagraph (3) (Il). 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 282 and 283 of this document.
153A. $ 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 153 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.
154. § 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).
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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155. § 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.
156. § 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.
264 Par 155
(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).
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.
The rotor operating speed limits shown during the official FAAAUTHORITY 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.
156A. § 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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157. § 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.
Inertia loads generated by linear and angular accelerations of the auxiliary rotors and their gear boxes, combined with
(il) 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.i 157A. § 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.
266 Par 157
158. § 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 (l.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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