126.-134. RESERVED.
Intentionally Left Blank 172 (thru 180) Par 126
135. RESERVED.
Intentionally Left Blank Par 135 181
SECTION 8. FLIGHT LOADS
136. § 27.321 (Amendment 27-11) GENERAL.
a. Explanation.
(1) The rule specifies the way the loads will be applied to the rotorcraft. It requires load analysis from minimum to maximum design weight. Any practical distribution of disposable loads must be included in the analysis.
(2) Paragraph (a) of the rule states:The flight load factor must be assumed to act normal to the longitudinal axis of the rotorcraft, and to be equal in magnitude and opposite in direction to the rotorcraft inertia load factor at the center of gravityi
b. Procedures.
(1) Derivation of the flight loads is required by and specified in §§ 27.337 through 27.351. This rule requires flight load determination from minimum to maximum weight and for disposable loads.
(2) The application of the design loads derived from the flight load factor will be as specified. The flight loads analysis data must comply with the rule.
137. § 27.337 (Amendment 27-26) LIMIT MANEUVERING LOAD FACTOR.
a. Explanation. The rotorcraft must be designed and substantiated to load factors as specified to provide a minimum level of structural integrity of the rotorcraft airframe and rotors.
(1) A range of design positive load factors from +3.5 to +2.0 may be used.
(2) A range of design negative load factors from -1.0 to -0.5 may be used.
(3) Load factors inside the range of +3.5 to -1.0 may be used provided the probability of exceeding the design load factors is shown by analysis and flight tests to be extremely remote and the selected load factors are appropriate to each weight condition between design maximum and minimum weight.
(4) Load factors exceeding these "minimums" may be used.
b. Procedures.
182 Par 136
(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.
The maximum positive design load factor of +3.5 is 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 § 27.321(b)(1) since the design load factor varies with gross weight.
(її) 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 § 27.337(b), the flight demonstration required by § 27.337 (b) must be conducted. The flight test personnel should determine that the demonstration shows 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.
Par 137 183
+3.5 Represents curve of Design thrust + lift Load Factor
• +2.0 Nz -1.0 Minimum Gross Weight Lbs. Maximum
LOAD FACTOR GROSS WEIGHT CURVE FIGURE 137-1 184 Par 137
138. § 27.339 (Amendment 27-11) RESULTANT LIMIT MANEUVERING LOADS.
a. Explanation. The rule specifies or defines the application of rotor and lift surtace loads to the rotorcraft.
(1) The design maneuvering load factors required by § 27.337 will result in or be derived from rotor thrust or lift and from auxiliary surface lift.
(2) Sections 27.321, 27.337, and 27.341 all complement one another and result in the derivation of design flight loads that will be imposed to ensure 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.
(Il) 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 $$ 27.321, 27.337, and 27.341. These loads are applied to the rotor or rotors and any auxiliary surface as prescribed by this rule.
139. §27.341 GUST LOADS.
a. Explanation.
(1) The rotorcraft must be substantiated for the loads derived from 30 feet per second vertical gusts from hovering to 1.11 VNE (i.e., VD).
Par 138 185
AC 27-1A 7130/97
(2) Gust loads for any horizontal stabilizing surface should be derived for vertical gusts, upward and downward.
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.
(ii) 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 is generally negligible when compared to the rotor and may be neglected for the rotorcraft gust load determination if proven negligible by analysis.
(Hi) The rotorcraft shall be assumed in stabilized level flight prior to meeting the gust.
(iv) The gust velocity may be assumed uniform across the rotorcraft.
(V) 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 ensure 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 Characteristics Including Unsteady Aerodynamics Stall Effects,i was written by P.J.
Arcidiacono, R.R. Berquist, and W.T. Alexander, Jr. References listed in the paper may be helpful also.
140. § 27.351 (Amendment 27-26) YAWING CONDITIONS.
a. Explanation. The standard was added by Amendment 27-26. It requires proof of a rotorcraft "structural" yaw or sideslip design envelope. This sideslip envelope 186 Par 139
should cover minimum forward speed or hover to the lesser of VH or VNE for "power-on" condition, not "power-off" since Vy is a part of the standard. The rotorcraft should be structurally safe for the thrust capability of the directional control system as stated.
(1) The rotorcraft structure should 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 derived from § 27.397(a)), is required for the conditions cited in the standard. A control system rate limiter or a yaw damper may be used as part of the type design required equipment, if elected. The total displacement is therefore a function of time as well as the maximum effort applied such as 130 pounds.
(i> At low airspeeds from 0 to 0.6 VNE, 90° yaw (sideward flight) is specified as the design limit.
(Ii) At high airspeeds VH or VNE), stabilized yaw angle (stabilized sideslip) or 15° sideslip, whichever is less, is specified to be substantiated.
(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 § 27.351(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.
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 Model FH1100. The Hiller FH1100 tail surface houses the tail rotor drive shaft and the tail rotor output gearbox.
(1) 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.
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NOTE: For one example: At stabilized yaw to the right (left pedal depressed to limit) (§ 27.351(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 § 27.351(b)(3), the tail rotor thrust moment then is added to the vertical stabilizer restoring moment. The addition of tail rotor thrust (§ 27.351(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 § 27.351(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 logical assessment of the various yaw conditions may be used to reduce the load derivation and analysis workload for critical rotorcraft design conditions.
(vi) The rotorcraft structure should be analyzed and/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 and 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 0.6 VNE) and 15° yaw when they are applied at the lesser of VNE or VH, with a straight line variation from 0.6 VNE to VH or VNE. The rotorcraft does not need to be capable of attaining the 90° and 15° yaw. They should be considered as the maximum sideslip limits.
188 Par 140
90°
YAW DEGREES (LONGITUDINAL AXIS) → 15° MAX.
0°
0.6 VNE AIRSPEED
LESSER OF
VA OR VNE FIGURE 140-1. SAMPLE YAW/FORWARD SPEED DIAGRAM Par 140 189
(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. However, this test is not a requirement.
(2) FAR § 27.351(b)(1) incorrectly references § 27.395(a) for the maximum pilot forces. The correct reference should be § 27.397(a).
141. § 27.361 (Amendment 27-23) ENGINE TORQUE.
a. Explanation.
(1) The rotorcraft shall be designed for limit engine torque values, as prescribed by the rule, to account for maximum engine torque, including certain transients and torsional oscillations. Amendment 27-23 separated the standard into paragraphs for turbine and reciprocating engine limit torque values.
(2) Turbine engine limit torque for design purposes (Amendment 27-23) was redefined into four cases and the torque values determined will be used. For example, sudden engine stoppage is introduced as one of the cases which is applied to the engine and the engine suspension and restraint system. Emergency operation of governor-controlled turboshaft engines is another case.
(3) Torque factors are also specified for reciprocating engines having two or more cylinders in Paragraph (b) of the standard.
(4) Sections 27.547(e)(1)(ii) and 27.549(d), respectively, refer to the application of engine torque to design of main rotor structure and engine mount and adjacent structure.
b. Procedures.
(1) The engine torque associated with the maximum continuous (MC) power condition for reciprocating engines should be multiplied by the appropriate torque factor to obtain the limit engine torque value used for structural substantiation of the rotorcraft.
(2) The torque values associated with MC power at 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.
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(3) Turbine engine limit torque values are determined for the four cases specified. Two cases are related to the endurance test of §§ 27.923 and 27.927.
(4) For sudden stoppage of turbine engines the engine manufacturers can reasonably provide FAA/AUTHORITY approved data to the applicant on inertia of rotating parts and the deceleration time expected in the event of sudden engine stoppage. This condition usually generates critical loads in the engine mounting and restraint system. These manufacturer's data should be acceptable for use in compliance with this part of the standard.
142.-151. RESERVED Par 141 191 (thru 201)