Sec. A23.1 General.

(a) The design load criteria in this Appendix are an approved equivalent of those in Secs. 23.321 through 23.399 of this subchapter for the certification of conventional, single-engine airplanes of 6,000 pounds [2700 kg] or less maximum weight.

(b) Unless otherwise stated, the nomenclature and symbols in this Appendix are the same as the corresponding nomenclature and symbols in Part 23.

Sec. A23.3 Special symbols.

*

*

*

*

* Also see paragraph A23.7(e)(2) of this Appendix.

Sec. A23.5 Certification in more than one category.

The criteria in this appendix may be used for certification in the normal, utility, and acrobatic categories, or in any combination of these categories. If certification in more than one category is desired, the design category weights must be selected to make the term

Sec. A23.7 Flight loads.

(a) Each flight load may be considered independent of altitude and, except for the local supporting structure for dead weight items, only the maximum design weight conditions must be investigated.

(b) Table 1 and figures 3 and 4 of this appendix must be used to determine values of

(c) Figures 1 and 2 of this appendix must be used to determine values of

(d) Each specified wing and tail loading is independent of the center of gravity range. The applicant, however, must select a c.g. range, and the basic fuselage structure must be investigated for the most adverse dead weight loading conditions for the c.g. range selected.

(e) The following loads and loading conditions are the minimums for which strength must be provided in the structure:

(1)

(2)

(3)

Sec. A23.9 Flight conditions.

(a)

(b)

(1) The airplane must be designed for at least the four basic flight conditions, "A", "D", "E", and "G" as noted on the flight envelope of figure 4 of this appendix. In addition, the following requirements apply:

(i) The design limit flight load factors corresponding to conditions "D" and "E" of figure 4 must be at least as great as those specified in Table 1 and figure 4 of this Appendix, and the design speed for these conditions must be at least equal to the value of

(ii) For conditions "A" and "G" of figure 4, the load factors must correspond to those specified in Table 1 of this Appendix, and the design speeds must be computed using these load factors with the maximum static lift coefficient determined by the applicant. However, in the absence of more precise computations, these latter conditions may be based on a value of =±1.35 and the design speed for condition "A" may be less than

(iii) Conditions "C" and "F" of figure 4 need only be investigated when

(2) If flaps or other high lift devices intended for use at the relatively low airspeed of approach, landing, and takeoff, are installed, the airplane must be designed for the two flight conditions corresponding to the values of limit flap-down factors specified in Table 1 of this appendix with the flaps fully extended at not less than the design flap speed

(c)

(1) The aft fuselage-to-wing attachment must be designed for the critical vertical surface load determined in accordance with subparagraphs A23.11(c)(1) and (2) of this Appendix.

(2) The wing and wing carry-through structures must be designed for 100 percent of condition "A" loading on one side of the plane of symmetry and 70 percent on the opposite side for certification in the normal and utility categories, or 60 percent on the opposite side for certification in the acrobatic category.

(3) The wing and wing carry-through structures must be designed for the loads resulting from a combination of 75 percent of the positive maneuvering wing loading on both sides of the plane of symmetry and the maximum wing torsion resulting from aileron displacement. The effect of aileron displacement on wing torsion at

(i) (up aileron side) wing basic airfoil.

(ii) (down aileron side) wing basic airfoil, where is the up aileron deflection and is the down aileron deflection.

(4) critical, which is the sum of + , must be computed as follows:

(i) Compute and from the formulas:

where = the maximum total deflection (sum of both aileron deflections) at V

(ii) Compute

where is the down aileron deflection corresponding to , and is the down aileron deflection corresponding to as computed in step (i).

(iii) If

(iv) If

(d)

(1) In designing the rear lift truss, the special condition specified in Sec. 23.369 may be investigated instead of condition "G" of figure 4 of this appendix. If this is done, and if certification in more than one category is desired, the value of

(2) Each engine mount and its supporting structures must be designed for the maximum limit torque corresponding to METO power and propeller speed acting simultaneously with the limit loads resulting from the maximum positive maneuvering flight load factor

(3) Each engine mount and its supporting structure must be designed for the loads resulting from a lateral limit load factor of not less than 1.47 for the normal and utility categories, or 2.0 for the acrobatic category.

Sec. A23.11 Control surface loads.

(a)

(b)

(c)

(1) Simplified limit surface loadings and distributions for the horizontal tail, vertical tail, aileron, wing flaps, and trim tabs are specified in Table 2 and figures 5 and 6 of this Appendix. If more than one distribution is given, each distribution must be investigated.

(2) If certification in the acrobatic category is desired, the horizontal tail must be investigated for an unsymmetrical load of 100 percent on one side of the airplane centerline and 50 percent on the other side of the airplane centerline.

(d)

(e)

Sec. A23.13 Control system loads.

(a)

(1) The flight control system and its supporting structure must be designed for loads corresponding to 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in A23.11 of this Appendix. In addition--

(i) The system limit loads need not exceed those that could be produced by the pilot and automatic devices operating the controls; and

(ii) The design must provide a rugged system for service use, including jamming, ground gusts, taxiing downwind, control inertia, and friction.

(2) Acceptable maximum and minimum limit pilot forces for elevator, aileron, and rudder controls are shown in the table in Sec. 23.397(b). These pilots loads must be assumed to act at the appropriate control grips or pads as they would under flight conditions, and to be reacted at the attachments of the control system to the control surface horn.

(b)

(c)

(d)

Table 1 - Limit Flight Load Factors

LIMIT FLIGHT LOAD FACTORS | |||||

Normal category | Utility category | Acrobatic category | |||

FLIGHT Load Factors | Flaps Up | n_{1} | 3.8 | 4.4 | 6.0 |

n_{2} | -0.5 n_{1} | ||||

n_{3} | Find n from Fig. 1_{3} | ||||

n_{4} | Find n_{4}_{ }from Fig. 2 | ||||

Flaps Down | n_{flap} | 0.5 n_{1} | |||

n_{flap} | Zero * |

1. Conditions "C" or "F" need only be investigated when or is greater than respectively.

2. Condition "G" need not be investigated when the supplementary condition specified in Sec. 23.369 is investigated.

Sec. B23.1 General.

(a) If allowed by the specific requirements in this part, the values of control surface loading in this appendix may be used to determine the detailed rational requirements of Secs. 23.397 through 23.459 unless the Administrator finds that these values result in unrealistic loads.

(b) For a seaplane version of a landplane, the landplane wing loadings may be used to determine the limit maneuvering control surface loadings (in accordance with B23.11 and figure 1 of Appendix B) if--

(1) The power of the seaplane engines does not exceed the power of the landplane engines:

(2) The placard maneuver speed of the seaplane does not exceed the placard maneuver speed of the landplane;

(3) The maximum weight of the seaplane does not exceed the maximum weight of the landplane by more than 10 percent;

(4) The landplane service experience does not show any serious control-surface load problem; and

(5) The landplane service experience is of sufficient scope to ascertain with reasonable accuracy that no serious control-surface load problem will develop on the seaplane.

Sec. B23.11 Control surface loads.

Acceptable values of limit average maneuvering control-surface loadings may be obtained from figure 1 of this Appendix in accordance with the following:

(a) For horizontal tail surfaces--

(1) With the conditions in Sec. 23.423(a), obtain as a function of

(i) Curve C of figure 1 for a deflection of 10° or less;

(ii) Curve B of figure 1 for a deflection of 20°;

(iii) Curve A for a deflection of 30° or more;

(iv) Interpolation for all other deflections; and

(v) The distribution of figure 7; and

(2) With the conditions in Sec. 23.423(b), obtain from curve B of figure (1) using the distribution of figure 7.

(b) For vertical tail surfaces--

(1) With the conditions in Sec. 23.441(a)(1), obtain as a function of

(2) With the conditions in Sec. 23.441 (a)(2), obtain from Curve C, using the distribution of figure 6; and

(3) With the conditions in Sec. 23.441 (a)(3), obtain from Curve A, using the distribution of figure 8.

(c) For ailerons, obtain from Curve B, acting in both the up and down directions, using the distribution of figure 9.

Figure 1 - Limit Average Maneuvering Control Surface Loading

Insert Table

INsert table

Insert table

Sec. C23.1 Basic landing conditions.

Tail wheel type | Nose wheel type | ||||

Condition | Level landing | Tail-down landing | Level landing with inclined reactions | Level landing with nose wheel just clear of ground | Tail-down landing |

Reference section | 23.479(a)(1) | 23.481(a)(1) | 23.479(a)(2)(i) | 23.479(a)(2)(ii) | 23.481(a)(2) and (b) |

Vertical component at c.g. | nW | nW | nW | nW | nW |

Fore and aft component at c.g. | KnW | 0 | KnW | KnW | 0 |

Lateral component in either direction at c.g. | 0 | 0 | 0 | 0 | 0 |

Shock absorber extension (hydraulic shock absorber) | Note (2) | Note (2) | Note (2) | Note (2) | Note (2) |

Shock absorber deflection (rubber or spring shock absorber) | 100% | 100% | 100% | 100% | 100% |

Tire deflection | Static | Static | Static | Static | Static |

Main wheel loads (both wheels)- | (n-L)W KnW | (n-L)Wb/d 0 | (n-L)Wa'/d' KnWa'/d' | (n-L)W KnW | (n-L)W 0 |

Tail (nose) wheel loads- | 0 0 | (n-L)Wa/d 0 | (n-L)Wb'/d' KnWb'/d' | 0 0 | 0 0 |

Notes | (1), (3), and (4) | -------(4)--------- | (1) | (1), (3), and (4) | (3) and (4) |

Note (2). For the purpose of design, the maximum load factor is assumed to occur throughout the shock absorber stroke from 25 percent deflection to 100 percent deflection unless otherwise shown and the load factor must be used with whatever shock absorber extension is most critical for each element of the landing gear.

Note (3). Unbalanced moments must be balanced by a rational or conservative method.

Note (4).