(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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Figure AC 29.427-1 (View Looking Forward)
АС 29-2C 9/30/99 Figure AC 29.427-2 (View Looking Forward) Page C -32
SUBPART C - STRENGTH REQUIREMENTS
GROUND LOADS AC 29.471. § 29.471 GROUND LOADS - GENERAL.
a. Explanation. This regulation specifies that limit ground loads must be considered which are:
(1) External loads caused by landing (ground) conditions and by ground taxiing loads as specified in § 29.235.
(2) Loads considering the rotorcraft structure as a rigid body.
(3) Loads in equilibrium with linear and angular inertia loads.
(4) The critical center of gravity “must be selected so that the maximum design loads are obtained in each landing gear element.”
b. Procedures.
(1) The standards to be considered are specified in §§ 29.473 through 29.511.
These associated standards cover landing gear arrangements, landing conditions, and ground handling conditions.
(2) Drop tests are required for determination of landing load factors. See paragraph AC 29.723.
(3) The application of the design loads derived from the landing load factors will be as specified for each element affected by landing or ground handling loads.
(4) During the applicant’s flight test program, the ground, landing, and taxiing load factors may be monitored to assure the design load factors used are adequate.
See paragraph AC 29.235 for § 29.235 guidance.
AC 29.473. § 29.473 (Amendment 29-3) GROUND LOADING CONDITIONS AND ASSUMPTIONS.
a. Explanation. The rotorcraft is to be designed for the maximum weight. A rotor lift of two-thirds of the design maximum weight may be used. The minimum limit landing load factor is determined by the drop tests of § 29.725. Provisions are made for supplementary energy absorption devices that have triggering mechanisms.
b. Procedures. Loads for the landing conditions are derived considering mass (equal to the maximum weight) and rotor lift (equal to two-thirds of the maximum weight) acting through the center of gravity throughout the landing impact. Unbalanced external Page C - 33
loads resulting from asymmetric loading conditions are reacted as specified in the individual subparagraphs.
NOTE: If supplementary energy absorption devices are used, neither they nor their triggering devices may fail under the loads established by the limit drop tests or the reserve energy absorption drop tests.
AC 29.475. § 29.475 TIRES AND SHOCK ABSORBERS.
a. Explanation. This section specifies the tire and shock absorber position to be used in ground load derivations.
b. Procedures. Ground loads are to be derived with the tires in static (1g) position and the shock absorbers “in their most critical position.” The determination of the “most critical position” for the shock absorbers generally requires a load versus deflection test or analysis of the shock absorber system and a determination of the effect of both load and deflections on the shock absorber, attachment structure, and substructure designed by ground loads.
AC 29.477. § 29.477 LANDING GEAR ARRANGEMENT.
a. Explanation. This section specifies the individual standards to be used for ground load conditions for rotorcraft having two wheels aft and one or more wheels forward of the center of gravity.
NOTE: § 29.497 gives ground loading conditions for landing gear with tail wheels, and § 29.501 gives ground loading conditions for landing gear with skids.
b. Procedures. The ground loading conditions of §§ 29.235, 29.479 through
29.485, and 29.493 will be used for rotorcraft having two wheels aft and one or more wheels forward of the center of gravity. This includes forward wheels on separa te axles.
AC 29.479. § 29.479 LEVEL LANDING CONDITIONS.
a. Explanation. This section provides explicit level landing load criteria for landing gear with two wheels aft and one or more wheels forward of the center of gravity.
(1) Level landings--
(i) Each wheel contacting the ground simultaneously; and
(ii) Aft wheels contacting the ground with forward wheels just clear of the ground.
(2) Application of loads-- Page C - 34
(i) Maximum design vertical loads applied alone;
(ii) The maximum design vertical loads applied with a drag load of at least 25 percent of the vertical load (applied at the ground contact area); and
(iii) The vertical load at the instant of peak drag load in conjunction with the peak drag load. A ground speed and load application is specified.
(3) A 40 percent/60 percent load distribution between wheels for configurations having two forward wheels including quadricycle. This distribution between wheels on a common axis is to be applied for the conditions of vertical loads only, and for vertical loads combined with drag loads of 25 percent of the vertical loads. Section 29.511 concerns a 60 percent to 40 percent ground load distribution between multiple-wheel units. See paragraph AC 29.511 for dual wheels on a common axle or axis.
(4) Aircraft pitching moments are to be reacted by the forward landing gear or by the angular inertia forces when the forward landing gear is clear of the ground as specified.
b. Procedures.
(1) The specified loading conditions will be used in load derivations.
(2) The critical center of gravity condition will be used for each gear and gear support structure.
(i) The aft center of gravity condition with the forward gear clear will normally be critical for the aft gear and gear supports.
(ii) The forward center of gravity condition with each gear contacting the ground simultaneously will normally design forward gear elements critical for vertical loads.
(iii) The forward center of gravity condition with the forward gear clear may result in high load factors, angular plus linear, that will greatly affect security of items of significant mass.
(3) The vertical load, at the instant of peak drag load combined with the peak drag component, can be determined from drop tests utilizing wheel spin-up or it can be analytically determined. If analysis is used, it must successfully correlate with the results of a previous well-instrumented test program.
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