Part V - Airworthiness Manual Chapter 523 - Normal, Utility, Aerobatic And Commuter Category Aeroplanes

Canadian Aviation Regulations (CARs) 2017-2

Content last revised: 2002/01/29

Preamble

SUBCHAPTERS 

  • A (523.1-523.3), 
  • B (523.21-523.253), 
  • C (523.301-523.575), 
  • D (523.601-523.871), 
  • E (523.901-523.1203), 
  • F (523.1301-523.1461), 
  • G (523.1501-523.1589)

APPENDICES 

ABCDEFGHIJ

(2002/03/01)

APPENDIX A

Simplified Design Load Criteria

A523.1 General

  1. (a) The design load criteria in this Appendix are an approved equivalent of those in 523.321 through 523.459 of this subchapter for an aeroplane having a maximum weight of 6,000 pounds or less and the following configuration:

    1. (1) A single engine excluding turbine powerplants; 

    2. (2) A main wing located closer to the aeroplane's centre of gravity than to the aft, fuselage-mounted, empennage;

    3. (3) A main wing that contains a quarter-chord sweep angle of not more than 15 degrees fore or aft;

    4. (4) A main wing that is equipped with trailing-edge controls (ailerons or flaps, or both);

    5. (5) A main wing aspect ratio not greater than 7;

    6. (6) A horizontal tail aspect ratio not greater than 4;

    7. (7) A horizontal tail volume coefficient not less than 0.34;

    8. (8) A vertical tail aspect ratio not greater than 2;

    9. (9) A vertical tail platform area not greater than 10 percent of the wing platform area; and

    10. (10) Symmetrical airfoils must be used in both the horizontal and vertical tail designs.

  2. (b) Appendix A criteria may not be used on any aeroplane configuration that contains any of the following design features:

    1. (1) Canard, tandem-wing, close-coupled, or tailless arrangements of the lifting surfaces;

    2. (2) Biplane or multiplane wing arrangements;

    3. (3) T-tail, V-tail, or cruciform-tail (+) arrangements;

    4. (4) Highly-swept wing platform (more than 15-degrees of sweep at the quarter-chord), delta planforms, or slatted lifting surfaces; or

    5. (5) Winglets or other wing tip devices, or outboard fins.

(Change 523-5)

A523.3 Special Symbols

n1 = Aeroplane Positive Manoeuvring Limit Load Factor.

n2 = Aeroplane Negative Manoeuvring Limit Load Factor.

n3 = Aeroplane Positive Gust Limit Load Factor at VC.

n4 = Aeroplane Negative Gust Limit Load Factor at VC.

nflap = Aeroplane Positive Limit Load Factor With Flaps Fully Extended at VF.

* VF min = Minimum Design Flap Speed =    kts.

* VA min = Minimum Design Manoeuvring Speed =    kts.

* VC min = Minimum Design Cruising Speed =   kts.

* VD min = Minimum Design Dive Speed =   kts.

* Also see paragraph A523.7(e)(2) of this Appendix.

A523.5 Approval in More Than One Category

The criteria in this Appendix may be used for certification in the normal, utility, and aerobatic 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", for the category corresponding to the maximum design weight, where "n1W" is constant.  If the aerobatic category is selected, a special unsymmetrical flight load investigation in accordance with subparagraphs A523.9(c)(2) and A523.11(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 aeroplane.  The engine mount, however, must be designed for a higher side load factor, if certification in the aerobatic category is desired, than that required for certification in the normal and utility categories.  When designing for landing loads, the landing gear and the aeroplane 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 Minister may require additional investigations for aircraft with unusual design features.

A523.7 Flight Loads

  1. (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.

  2. (b) Table Al and Figures A3 and A4 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.

  3. (c) Figures Al and A2 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.

  4. (d) Each specified wing and tail loading is independent of the centre 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.

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

    1. (1) Aeroplane 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 aeroplane normal loads (as determined from paragraphs A523.9(b) and (c) of this Appendix) for the positive flight conditions and a magnitude equal to the aeroplane normal loads for the negative conditions.  Each chordwise and normal component of this wing load must be considered.

  1. (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 A3  of this Appendix.  In addition, VCmin need not exceed values of 0.9 VH 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.

  2. (3) Flight load factor.  The limit flight load factors specified in Table Al of this Appendix represent the ratio of the aerodynamic force component (acting normal to the assumed longitudinal axis of the aeroplane) to the weight of the aeroplane.  A positive flight load factor is an aerodynamic force acting upward, with respect to the aeroplane.

A523.9 Flight Conditions

  1. (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 aeroplane similar to the diagram in Figure A4 of this Appendix.  This diagram must also be used to determine the aeroplane structural operating limitations as specified in 523.1505 through 523.1511 and 523.1519.

  1. (b) Symmetrical flight conditions. The aeroplane must be designed for symmetrical flight conditions as follows:

    1. (1) The aeroplane must be designed for at least the four basic flight conditions, "A", "D", "E", and "G" as noted on the flight envelope of Figure A4 of this Appendix.  In addition, the following requirements apply:

      1. (i) The design limit flight load factors corresponding to conditions "D" and "E" of Figure A4 must be at least as great as those specified in Table Al and Figure A4 of this Appendix, and the design speed for these conditions must be at least equal to the value of VD found from Figure A3  of this Appendix.

      2. (ii) For conditions "A" and "G" of Figure A4, the load factors must correspond to those specified in Table Al of this Appendix, and the design speeds must be computed using these load factors with the maximum static lift coefficient CNA determined by the applicant.  However, in the absence of more precise computations, these latter conditions may be based on a value of CNA = +1.35 and the design speed for condition "A" may be less than VAmin.

      3. (iii) Conditions "C" and "F" of Figure A4 need only be investigated when n3W/S or n4W/S are greater than n1W/S or n2W/S of this Appendix, respectively.

    2. (2) If flaps or other high lift devices intended for use at the relatively low airspeed of approach, landing, and takeoff, are installed, the aeroplane must be designed for the two flight conditions corresponding to the values of limit flap-down factors specified in Table Al of this Appendix with the flaps fully extended at not less than the design flap speed VFmin from Figure A3 of this Appendix.

  1. (c) Unsymmetrical flight conditions.  Each affected structure must be designed for unsymmetrical loadings as follows:

    1. (1) The aft fuselage-to-wing attachment must be designed for the critical vertical surface load determined in accordance with subparagraphs A523.11(c)(1) and (2) of this Appendix.

  1. (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 aerobatic category.

  2. (3) The wing and wing carry-through structures must be designed for the loads resulting from a combination of 75 percent of the positive manoeuvring 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:

    1. (i) Cm = Cm + 0.01du (up aileron side) wing basic airfoil.

    2. (ii) Cm = Cm - 0.01dd (down aileron side) wing basic airfoil, where du is the up aileron deflection and dd is the down aileron deflection.

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

    1. (i) Compute  a and  b from the formulas:

  and

where:   p = the maximum total deflection (sum of both aileron deflections) at VA with VA, VC, and VD described in subparagraph (2) of A523.7 (e) of this Appendix.

(ii) Compute K from the formula:

where da is the down aileron deflection corresponding to  a and db is the down aileron deflection corresponding to  b as computed in step (i).

(iii) If K is less than 1.0,  a is   critical and must be used to determine du and dd.  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 du and dd. 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 523.369 may be investigated instead of condition "G" of Figure A4 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 523.369 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 manoeuvring 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 aerobatic category.

A523.11 Control Surface Loads

  1. (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.

  1. (b) Limit pilot forces.  In each control surface loading condition described in paragraphs (c) through (e) of this section, the air loads 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 523.397(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 A2 of this Appendix.

  1. (c) Surface loading conditions.  Each surface loading condition must be investigated as follows:

  1. (1) Simplified limit surface loadings for the horizontal tail, vertical tail, aileron, wing flaps, and trim tabs are specified in Figures A5 and A6 of this Appendix.

    1. (i) The distribution of load along the span of the surface, irrespective of the chord-wise load distribution, must be assumed proportional to the total chord, except on horn balance surfaces.

    2. (ii) The load on the stabilizer and elevator, and the load on fin and rudder, must be distributed chord-wise as shown in Figure A7 of this Appendix.

    3. (iii) In order to ensure adequate torsional strength and to account for manoeuvres and gusts, the most severe loads must be considered in association with every centre of pressure position between the leading edge and the half chord of the mean chord of the surface (stabilizer and elevator, or fin and rudder).

    4. (iv) To ensure adequate strength under high leading edge loads, the most severe stabilizer and fin loads must be further considered as being increased by 50 percent over the leading 10 percent of the chord with the loads aft of this appropriately decreased to retain the same total load.

    5. (v) The most severe elevator and rudder loads should be further considered as being distributed parabolically from three times the mean loading of the surface (stabilizer and elevator, or fin and rudder) at the leading edge of the elevator and rudder, respectively, to zero at the trailing edge according to the equation:

Figure

Where:

P(x) = local pressure at the chord-wise stations x,

c = chord length of the tail surface,

cf = chord length of the elevator and rudder respectively, and

  = average surface loading as specified in Figure A5.

  1. (vi) The chord-wise loading distribution for ailerons, wing flaps, and trim tabs are specified in Table A2 of this Appendix.

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

  1. (d) Outboard fins.  Outboard fins must meet the requirements of 523.445.

  2. (e) Special devices.  Special devices must meet the requirements of 523.459.

(Change 523-5)

A523.13 Control System Loads

  1. (a) Primary flight controls and systems.  Each primary flight control and system must be designed as follows:

    1. (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 A523.11 of this Appendix.  In addition:

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

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

    2. (2) Acceptable maximum and minimum limit pilot forces for elevator, aileron, and rudder controls are shown in the table in 523.397(b).  These pilot 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.

  2. (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 pilot forces shown in the table in 523.397(b).

  3. (c) Ground gust conditions.  Ground gust conditions must meet the requirements of 523.415.

  4. (d) Secondary controls and systems.  Secondary controls and systems must meet the requirements of 523.405.

  • Table Al - Limit Flight Load Factors

  • Table A2 - Average Limit Control Surface Loading

  • Figure A-1 -  Chart for finding n3 factor at speed Vc

  • Figure A-2 - Chart for finding n4factor at speed Vc

  • Figure A3 - Determination of minimum design speeds equations [Speeds are in knots]

  • Figure A4 - Flight envelope

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

2. Condition "G" need not be investigated when the supplementary condition specified in 523.369 is investigated.

  • Figure A5 - Average limit control surface loading

  • Figure A6 - Average limit control surface loading

  • Figure A7 - Chordwise load distribution for Stabilizer and Elevator or fin and rudder

where:

  = average surface loading (as specified in Figure A5)

E = ratio of elevator (or rudder) chord to total stabilizer and elevator (or fin and rudder) chord.

d'= ratio of distance of centre of pressure of a unit span wise length of combined stabilizer and elevator (or fin and rudder) measured from stabilizer (or fin) leading edge to the local chord.  Sign convention is positive when centre of pressure is behind leading edge.

c = local chord.

Note:  Positive values of  , P1 and P2 are all measured in the same direction.

(Change 523-5)

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