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
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.
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
(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 category||Utility category||Acrobatic category|
Find n3 from Fig. 1
Find n4 from Fig. 2
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
(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
Sec. C23.1 Basic landing conditions.
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.
Tail wheel type
Nose wheel type
Level landing with inclined reactions
Level landing with nose wheel just clear of ground
|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.|
|Fore and aft component at c.g.|
|Lateral component in either direction at c.g.|
|Shock absorber extension (hydraulic shock absorber)|
|Shock absorber deflection (rubber or spring shock absorber)|
|Main wheel loads (both wheels)-|
|Tail (nose) wheel loads-|
|Notes||(1), (3), and (4)||-------(4)---------|
|(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). L is defined in Sec. 23.725(b).