FAR 23 Appendix A--Simplified Design Load Criteria for Conventional, Single-Engine Airplane of 6,000 Pounds or Less Maximum Weight

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.

n1 = Airplane Positive Maneuvering Limit Load Factor.
n2 = Airplane Negative Maneuvering Limit Load Factor.
n3 = Airplane Positive 30 fps Gust Limit Load Factor at VC.
n4 = Airplane Negative 30 fps Gust Limit Load Factor at VC.
nflap = Airplane Positive Limit Load Factor with Flaps Fully Extended at VF.
*VF min = Minimum Design Flap Speed =
*VA min = Minimum Design Maneuvering Speed =
*VC min = Minimum Design Cruising Speed =
*VD min = Minimum Design Dive Speed =
* 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 "n1W' constant for all categories or greater for one desired category than for others. The wings and control surfaces (including wing flaps and tabs) need only be investigated for the maximum value of "n1W', or for the category corresponding to the maximum design weight, where "n1W' is constant. If the acrobatic category is selected, a special unsymmetrical flight load investigation in accordance with subparagraphs A23.9(c)(2) and A23.11(c)(2) of this appendix must be completed. The wing, wing carry through and the horizontal tail structures must be checked for this condition. The basic fuselage structure need only be investigated for the highest load factor design category selected. The local supporting structure for dead weight items need only be designed for the highest load factor imposed when the particular items are installed in the airplane. The engine mount, however, must be designed for a higher side load factor, if certification in the acrobatic category is desired, than that required for certification in the normal and utility categories. When designing for landing loads, the landing gear and the airplane as a whole need only be investigated for the category corresponding to the maximum design weight. These simplifications apply to single-engine aircraft of conventional types for which experience is available, and the Administrator may require additional investigations for aircraft with unusual design features.


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 n1, n2,
n3, and n4, corresponding to the maximum design weights in the desired categories.
(c) Figures 1 and 2 of this appendix must be used to determine values of n3 and n4 corresponding to the minimum flying weights in the desired categories, and, if these load factors are greater than the load factors at the design weight, the supporting structure for dead weight items must be substantiated for the resulting higher load factors.
(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) Airplane equilibrium. The aerodynamic wing loads may be considered to act normal to the relative wind, and to have a magnitude of 1.05 times the airplane normal loads (as determined from paragraphs A23.9(b) and (c) of this appendix) for the positive flight conditions and a magnitude equal to the airplane normal loads for the negative conditions. Each chordwise and normal component of this wing load must be considered.
(2) Minimum design airspeeds. The minimum design airspeeds may be chosen by the applicant except that they may not be less than the minimum speeds found by using figure 3 of this appendix. In addition, VC min need not exceed values of 0.9VH actually obtained at sea level for the lowest design weight category for which certification is desired. In computing these minimum design airspeeds, n1 may not be less than 3.8.
(3) Flight load factor. The limit flight load factors specified in Table 1 of this appendix represent the ratio of the aerodynamic force component (acting normal to the assumed longitudinal axis of the airplane) to the weight of the airplane. A positive flight load factor is an aerodynamic force acting upward, with respect to the airplane.

Sec. A23.9 Flight conditions.

(a) General. Each design condition in paragraphs (b) and (c) of this section must be used to assure sufficient strength for each condition of speed and load factor on or within the boundary of a V-n diagram for the airplane similar to the diagram in figure 4 of this appendix. This diagram must also be used to determine the airplane structural operating limitations as specified in Secs. 23.1501(c) through 23.1513 and 23.1519.
(b) Symmetrical flight conditions. The airplane must be designed for symmetrical flight conditions as follows:
(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 VD found from figure 3 of this appendix.
(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 VA min.
(iii) Conditions "C" and "F" of figure 4 need only be investigated when n3 W/S or n4 W/S are greater than n1 W/S or n2 W/S of this appendix, respectively.
(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 VF min from figure 3 of this appendix.
(c) Unsymmetrical flight conditions. Each affected structure must be designed for unsymmetrical loadings as follows:
(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 VC or VA using the basic airfoil moment coefficient modified over the aileron portion of the span, must be computed as follows:
(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 VA with VA, VC, and VD described in subparagraph (2) of Sec. 23.7(e) of this appendix.
(ii) Compute K from the formula:

where is the down aileron deflection corresponding to , and is the down aileron deflection corresponding to as computed in step (i).
(iii) If K is less than 1.0, is critical and must be used to determine and . In this case, VC is the critical speed which must be used in computing the wing torsion loads over the aileron span.
(iv) If K is equal to or greater than 1.0, is critical and must be used to determine and . In this case, VD is the critical speed which must be used in computing the wing torsion loads over the aileron span.
(d) Supplementary conditions; rear lift truss; engine torque; side load on engine mount. Each of the following supplementary conditions must be investigated:
(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 W/S used in the formula appearing in Sec. 23.369 must be that for the category corresponding to the maximum gross weight.
(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 n1. The limit torque must be obtained by multiplying the mean torque by a factor of 1.33 for engines with five or more cylinders. For 4, 3, and 2 cylinder engines, the factor must be 2, 3, and 4, respectively.
(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) General. Each control surface load must be determined using the criteria of paragraph (b) of this section and must lie within the simplified loadings of paragraph (c) of this section.
(b) Limit pilot forces. In each control surface loading condition described in paragraphs (c) through (e) of this section, the airloads on the movable surfaces and the corresponding deflections need not exceed those which could be obtained in flight by employing the maximum limit pilot forces specified in the table in Sec. 23.397(b). If the surface loads are limited by these maximum limit pilot forces, the tabs must either be considered to be deflected to their maximum travel in the direction which would assist the pilot or the deflection must correspond to the maximum degree of "out of trim" expected at the speed for the condition under consideration. The tab load, however, need not exceed the value specified in Table 2 of this Appendix.
(c) Surface loading conditions. Each surface loading condition must be investigated as follows:
(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) Outboard fins. Outboard fins must meet the requirements of Sec. 23.455.
(e) Special devices. Special devices must meet the requirements of Sec. 23.459.

Sec. A23.13 Control system loads.

(a) Primary flight controls and systems. Each primary flight control and system must be designed as follows:
(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) Dual controls. If there are dual controls, the systems must be designed for pilots operating in opposition, using individual pilot loads equal to 75 percent of those obtained in accordance with paragraph (a) of this section, except that individual pilot loads may not be less than the minimum limit pilot forces shown in the table in Sec. 23.397(b).
(c) Ground gust conditions. Ground gust conditions must meet the requirements of Sec. 23.415.
(d) Secondary controls and systems. Secondary controls and systems must meet the requirements of Sec. 23.405.

Table 1 - Limit Flight Load Factors

LIMIT FLIGHT LOAD FACTORS
Normal categoryUtility categoryAcrobatic category
FLIGHT
Load Factors
Flaps Up
n1
3.8
4.4
6.0
n2
-0.5n1
n3
Find n3 from Fig. 1
n4
Find n4 from Fig. 2
Flaps Down
nflap
0.5n1
nflap
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.





Appendix B--Control Surface Loadings

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 W/S and surface deflection, using--
(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 W/S and surface deflection using the same requirements as used in subdivisions (a)(1)(i) through (a)(1)(v);
(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 section23.479(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 (1). K may be determined as follows: K = 0.25 for W = 3,000 pounds or less; K = 0.33 for W = 6,000 pounds or greater, with linear variation of K between these weights.
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). L is defined in Sec. 23.725(b).