(1) The applicant may elect to substantiate the rotorcraft for a design maneuvering load factor less than +3.5 and more than -1.0. Whenever this option is used, an analytical study and flight demonstration are required.
(I) The maximum positive design load factor is +3.5 generally at a weight below maximum gross weight. The maximum thrust capability of the main rotor combined with incremental lift of wings or sponsons, if installed, results in a maximum design positive load factor. An example of a load factor - gross weight curve is shown in Figure 137-1. Note the minimum positive design load factor is +2.0 even though the required analysis and flight demonstration may prove the rotorcraft is not capable of achieving this load factor. This curve also illustrates compliance with § 29.337(b)(2) since the design load factor varies with gross weight.
(ii) The largest negative design load factor is -1.0; however, several current rotorcraft designs are not capable of achieving a negative load factor.
Therefore, -0.5 has been an acceptable structural design negative load factor for certain rotorcraft designs.
(2) Whenever the applicant analytically substantiates the lower load factors allowed by § 29.337(b), the applicant must conduct the flight demonstration required by § 29.337(b)(1). The flight test personnel must determine that the demonstration is conducted in a manner to show that the probability of exceeding the selected design load factors, (those factors less than +3.5 and more than -1.0) is extremely remote.
(See Order 8110.4, Paragraph 166c(2) (c)).
(3) A numerical value has not been assigned to "extremely remote" in this standard.
240 Par 137
+3.5 Represents curve of Design thrust + lift Load Factor Nz +2.0
- 1.0 Minimum Gross Weight Lbs. Maximum
LOAD FACTOR GROSS WEIGHT CURVE FIGURE 137-1 Par 137 241
138. § 29.339 RESULTANT LIMIT MANEUVERING LOADS.
a. Explanation. The rule specifies or defines the application of rotor and lift surface loads to the rotorcraft.
(1) The design maneuvering load factors required by § 29.337 will result in or be derived from rotor thrust or lift and from auxiliary surface lift.
(L) The rules 99 29.321, 29.337, 29.341, and 29.351 all complement one another and result in the derivation of design flight loads that will be imposed to assure structural integrity of the rotorcraft.
(3) The following assumptions and conditions are specified in the rule.
The rule requires application of appropriate loads at each rotor hub and auxiliary lifting surface.
(ii) Power-on and power-off flight with maximum design rotor tip speed ratio and specific conditions that must be considered.
Rotor tip speed ratio, defined in the rule, has been carried forward from the initial rotorcraft certification rules issued in 1946. The rotor tip speed ratio is a basic parameter used in calculating rotor aerodynamic forces.
b. Procedures.
(1) The rule specifies an acceptable assumption concerning application of the rotorcraft maneuvering loads.
(2) The rotor tip speed ratio is a parameter found in textbooks and other books such as NACA Report No. 716. The equation in the rule contains angle, "a." Report No. 716 also defines angle, "a," as the angle of attack of the rotor disk. This definition is more easily understood than the definition contained in the rule.
(3) The rotorcraft design loads are derived as prescribed by §§ 29.321, 29.337,
29.341, and 29.351. These loads are applied to the rotor or rotors and any auxiliary surface as prescribed by this rule.
139. §29.341 GUST LOADS
a. Explanation.
(1) The rotorcraft must be substantiated for the loads derived from 30 feet per second vertical and horizontal gusts from hovering to 1.11 VNE: i.e., (VD).
242 Par 138
(2) Gust loads for any vertical stabilizing surface should be derived for lateral or sideward gusts, as well as the head-on horizontal gusts. See Paragraph No. 159, § 29.413(a)(2) of this advisory circular.
(3) Gust loads for any horizontal stabilizing surface should be derived for vertical gusts, upward and downward, as well as for head-on gusts. See Paragraph 159.
b. Procedures.
(1) Either sharp-edged (instantaneous) gusts or sharp-edged gusts modified by an alleviation (attenuation) factor may be used for calculating aerodynamic loads for the rotorcraft and any installed stabilizing surfaces. The following conditions may be used:
Vertical gusts may be considered normal to the flight path of the rotorcraft except during hover or low speed flight (20 knots or less) when the gusts may be assumed normal to the longitudinal axis of the rotorcraft.
(il) For a vertical stabilizing surface, the horizontal gusts are normal to the flight path of the rotorcraft except during hover or low speed flight when the gusts may be assumed normal to the longitudinal axis of the rotorcraft.
(III) A primary effect of encountering the gust is to change the lift of the rotors and rotorcraft surfaces. Of primary concern is the gust load or lift created by the main rotor or rotors. The lift increment of the horizontal stabilizing surface and fuselage are generally negligible when compared to the rotor and may be neglected for the rotorcraft gust load determination if proven negligible by analysis.
(iv) The rotorcraft shall be assumed in stabilized level flight prior to meeting the gust.
(v) The gust velocity may be assumed uniform across the rotorcraft.
(vi) Gust loads on the stabilizing surfaces are required as stated in Paragraph 159 of this advisory circular.
(2) The rotorcraft design maneuvering load factors may generally exceed the design gust load factors calculated in compliance with this rule. This may be attributed to the small incremental change in lift due to the 30 FPS gust. Nonetheless, design gust loads for the rotorcraft shall be calculated as specified in the rule to assure the rotorcraft maneuvering load factors do, in each case, exceed the design gust load tactor.
(3) For further information about rotorcraft gust response characteristics, see Paper No. 9 presented at the AHS/NASA - Ames Specialist's Meeting on Rotorcraft Par 139 243
Dynamics, February 13-15, 1974. The paper, entitled, "Helicopter Gust Response Characteristics Including Unsteady Aerodynamics Stall Effects," was written by P.J.
Arcidiacono, R.R. Berquist, and W.T. Alexander, Jr. References listed in the paper may be helpful also.
140. § 29.351 YAWING CONDITIONS.
a. Explanation. The rule requires proof of a rotorcraft "structural" yaw or sideslip design envelope. This sideslip envelope must cover minimum forward speed or hover to Vu or VNE, whichever is less. The rotorcraft must be structurally safe for the thrust capability of the directional control system.
(1) The rotorcraft structure must be designed to withstand the loads for the specified yaw conditions. The standard does not require a structural flight demonstration. It is a structural design standard.
(2) Maximum displacement of the directional control, except as limited by pilot effort (130 pounds; § 29.397(a)), is required for the conditions cited in the rule. A control system rate limiter or a yaw damper may be used. The total displacement is therefore a function of time as well as the maximum effort applied (130 pounds).
(i) At low airspeeds, 90° yaw (sideward flight) should be the design limit.
(ii) At high airspeeds, stabilized yaw angle (stabilized sideslip) must be substantiated as stated in the rule.
(ill) At high airspeeds, the maximum tail rotor thrust will be combined with the vertical (directional) stabilizer surface load, if a stabilizer is used, as specified by § 29.351(b)(1).
(iv) At high airspeeds, while the rotorcraft is in the sideslip condition, the directional control is then returned to the neutral position, attendant with the flight condition. The tail rotor thrust will be added to the restoring force of the vertical stabilizer.
(v) Both right and left yaw conditions should be proven.
(3) The tail rotor attachment structure must comply with § 29.403.
(4) The vertical stabilizing surface must also comply with § 29.413.
b. Procedures.
(1) Many of the current single main rotor rotorcraft designs have vertical (directional) stabilizing surfaces. These surfaces may be solely vertical stabilizing fins 244 Par 139
as on the Bell Model 206, or a swept vertical extension of the tail boom as on the Hiller Model FH1100. The Hiller FH1100 tail surface houses the tail rotor drive shaft and the tall rotor output gearbox.
For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust.
(її) For vertical stabilizers that house the tail rotor output gearbox, such as the Hiller Model FH1100, the tail surface air loads will add to or subtract from the tail rotor thrust according to the flight condition under consideration.
NOTE: For one example: At stabilized yaw to the right (left pedal depressed to limit) (§ 29.351(b)(2)), the tail rotor thrust moment should equal the restoring moment of the tail boom, vertical stabilizer and main rotor torque. As stated by § 29.351(b)(3), the tail rotor thrust moment then is added to the vertical stabilizer restoring moment. The addition of tail rotor thrust (§ 29.351(b)(3)) and vertical stabilizer load is generally one of the critical design conditions for the fuselage/tail boom.
(iii) For vertical stabilizers or fins that have an offset incidence angle with respect to the rotorcraft axis, the vertical fin moment is added, or subtracted as applicable, to the tail rotor thrust moment. The condition stated in § 29.351(b)(1) may result in adding the fin load to the tail rotor thrust.
(iv) Low airspeed maneuvers, such as sideward, rearward, and hover turns over a spot, typically impose insignificant aerodynamic loads on the fuselage and/or tail boom. The aerodynamic loads at VH or VNE, whichever is required, are generally the significant aerodynamic design loads.
(v) A rational assessment of the various yaw conditions may be used to reduce the load deviation and analysis to the critical rotorcraft design conditions.
(vi) The rotorcraft structure shall be analyzed or tested for loads derived from the critical design conditions.
(vii) A simple structural design envelope may be derived from these design data. If the right or left yaw limits are not very different, common, conservative design limits may be used. A sample yaw/forward speed diagram, as derived from design analysis of the characteristics of a hypothetical rotorcraft, is presented in Figure 140-1. A table of values would also suffice. This figure reflects characteristics which include a 90° yaw when the directional control inputs are applied at low airspeeds (up to 30 knots presumably the maximum sideward flight speed of which this aircraft is capable) and 10° yaw when they are applied at VH, with a straight line variation from 30 knot forward speed to VH.
Par 140 245
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).
Par 140 247
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.
Par 140B 249
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.
Par 152 261