+3.5 Represents curve of thrust + lift Design Load Factor Nz -+2.0
- 1.0 - Minimum Gross Weight Lbs. Maximum
LOAD FACTOR GROSS WEIGHT CURVE FIGURE AC 29.337-1 Page C - 7
AC 29.339. § 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.
(2) The rules §§ 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.
(i) 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.
(iii) 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.
AC 29.341. §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).
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(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 AC 29.413, § 29.413(a)(2).
(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 AC 29.413.
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:
(i) 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.
(ii) 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 AC 29.413.
(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 factor.
(3) For further information about rotorcraft gust response characteristics, see Paper No. 9 presented at the AHS/NASA - Ames Specialist’s Meeting on Rotorcraft Dynamics, February 13-15, 1974. The paper, entitled, “Helicopter Gust Response Page C - 9
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.
AC 29.351. § 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 VH 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.
(iii) 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 as on the Bell Model 206, or a swept vertical extension of the tail boom as on the Hiller Page C - 10
Model FH1100. The Hiller FH1100 tail surface houses the tail rotor drive shaft and the tail rotor output gearbox.
(i) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust.
(ii) 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 V H 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 AC 29.351-1. 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.
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- 90° 90° YAW DEGREES / 10° MAX 0°
AIRSPEED 30 KNOT LESSER OF
SIDEWARD VH OR VNE FLIGHT FIGURE AC 29.351-1 SAMPLE YAW/FORWARD SPEED DIAGRAM Page C - 12
(viii) 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.
AC 29.351A. § 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 VH 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 VH (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 AC 29.351A-1.
(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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go•
YAW DEGREES (LON GI TUDI N AL AXIS)
AIRSPEED +-15° MAX.
LESSER OF
VH OR VNE AC 29-2C, Chg 1 4/25/06 FIGURE AC 29.351A-1
SAMPLE YAW/FORWARD SPEED DIAGRAM Page C - 14
AC 29.351B. § 29.351 (Amendment 29-40) YAWING CONDITIONS.
a. Definitions.
(1) Suddenly. For the purpose of this section, ‘suddenly’ is defined as an interval not to exceed 0.2 seconds for complete control input. A rational analysis may be used to substantiate an alternative value.
(2) Zero Yaw. Normal, 1 -g, level flight condition with either zero bank angle or zero sideslip.
b. Explanation. The rule requires a rotorcraft “structural” yaw or sideslip design envelope. This sideslip envelope must cover minimum forward speed, or hover, to VH 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 yawing conditions. The standard does not require a structural flight demonstration. It is a structural design standard.
(2) This standard applies only to power-on conditions. Autorotations need not be considered.
(3) This standard requires the maximum allowable rotor RPM consistent with the flight conditions, including special operational rotor settings.
(4) For the purposes of this section, the analysis may be performed at international standard atmosphere (ISA) sea level conditions.
(5) The rotorcraft structure must be designed to withstand the loads for the specified sideslip conditions. This includes, but is not limited to:
(i) Main cabin, tailboom, and vertical control surfaces.
(ii) Tai l rotor structures, including the fitting attachments to the frame.
(iii) Windows, doors, and other transparencies.
(iv) Landing gear and retracting mechanism.
(v) Fairings and cowlings.
(6) Maxim um displacement of the directional control, except as limited by pilot effort (§ 29.397(a)), is required for the conditions cited in the rule. Control system limiting devices may be used, however the probability of failure or malfunction of these system(s) should be considered (see Figure AC 29.351B -2). This evaluation may Page C - 15
AC 29-2C, Chg 2 4/25/06 include Flight Manual Limitations, if failure of the system is reliably indicated to the crew.
(7) Both right and left yaw conditions should be evaluated.
(8) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust (superpositioning).
(9) Loads associated with sideslip angles exceeding the values of Figure AC 29.351B-1 do not need to be considered. The corresponding points of the maneuver may be deleted.
c. Procedure. The design loads should be evaluated within the limits of Figure AC 29.351B-1 or the maximum capability of the rotorcraft, whichever is less; at speeds from zero to VH or VNE, whichever is less, for the following phases of the maneuver:
(1) With the rotorcraft at an initial trim condition (1 g level flight and zero yaw), the cockpit directional control is suddenly displaced to the maximum deflection limited by the control stops or by the maximum pilot force specified in § 29.397(a). This is intended to generate a high tail rotor thrust.
(2) While maintaining maximum cockpit directional control deflection, within the limitation specified in c(1) of this AC paragraph, allow the rotorcraft to yaw to the maximum transient sideslip angle or to the value defined in Figure AC 29.351B-1, whichever is less. This is intended to generate high aerodynamic loads.
(3) Allow the rotorcraft to stabilize at the maximum steady-state sideslip angle.
In the event that the maximum steady state angle is greater than the value defined in Figure AC 29.351B-1, the rotorcraft should be trimmed to the value of the angle using less than maximum cockpit directional control deflection.
(4) With the rotorcraft yawed to the static equilibrium sideslip angle specified in c(3) of this AC paragraph, the cockpit control is suddenly returned to its initial trim position. This is intended to combine a high tail rotor thrust and high aerodyna mic restoring forces.
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SIDESLIP ANGLE ( deg. ) 90 15 0 0.6 Vne Vh or Vne whichever is less
AIRSPEED FIGURE AC 29.351B-1
YAW/FORWARD SPEED DIAGRAM Page C - 17
FS 1 .5 1 .0 -~ I 10-9 Qj - Probability of being in failure condition j AC 29-2C, Chg 2 4/25/06
• For static strength substantiation, each part of the structure should be able to withstand without failure, the loads generated by the maneuver described in the rule multiplied by a factor of safety depending on the probability of being in this failure state. The factor of safety is defined in the figure below:
Qj = (Tj)(Pj) where:
Tj = Average flight time spent with a failed control limiting system j (in hours) Pj = Probability of occurrence of failure of the control limiting system j (per hour) Note:If Pj is greater than 10-3 per flight hour then a 1.5 factor of safety should be applied to all limit load conditions specified in this standard.
FIGURE AC 29.351B-2 Safety Factors for Probability of Failure AC 29.361. § 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.
(i) 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.”
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