AC 43.13-2A

Rev. 1977 AC 43.13-2A Chapter 7. ROTORCRAFT EXTERNAL-LOAD DEVICE INSTALLATIONS Section 1. CARGO SLINGS

96. GENERAL. This section contains structural and design information for the fabrication and installation of a cargo sling used as an externalload attaching means for a Class B rotorcraftload combination operation under FAR Part 133.

As an external-load attaching means, a cargo sling" includes a quick-release device and the associated cables, fittings, etc., used for the attachment of the cargo sling to the rotorcraft.

97. QUICK-RELEASE DEVICE. Section 133.43 (d) of the FARs specifies the requirements for the quick-release device. In addition to commercially manufactured helicopter cargo hooks, some surplus military bomb releases meet the requirements of that section.

98. LOCATION OF CARGO RELEASE IN RELATION TO THE ROTORCRAFT'S C.G. LIMITS.

a. An ideal location of the cargo release would be one that allows the line of action to pass through the rotorcraft's center of gravity at all times. (See fig. 7.1, illus. A.) However, with « Lines of + Line of action action Effective pivot point Moment

• Arms Arm action Effective pivot Effective pivot pointe F F FIGURE 7.1.-Location of cargo release in relation to the rotorcraft's center of gravity.

Chap 7

AC 43.13-2A Rev. 1977 most cargo sling installations, this ideal situation is realized only when the line of action is vertical or near vertical and through the rotorcraft's c.g.

(See fig. 7.1, illus. B.)

b. Whenever the line of action does not pass through the rotorcraft's c.g. due to the attachment method used, acceleration forces, or aerodynamic forces, the rotorcraft-load combined center of gravity will shift from the rotorcraft's c.g. position. Depending upon the factors involved, the shift may occur along either or both the longitudinal or lateral axes. The amount of shift is dependent upon the force applied (F) and the length of the arm of the line of action. Their product (F x Arm) yields a moment which can be used to determine the rotorcraft-load combined center of gravity. (See fig. 7. 1, illus. C.) If the rotorcraft-load center of gravity is allowed to shift beyond the rotorcraft's approved center of gravity limits, the rotorcraft may become violently uncontrollable.

c. Thus, any attachment method or location which will decrease the length of the arm will reduce the distance that the combined center of gravity will shift for a given load (F) and line of action angle. (See fig. 7.1, illus. D.)

99. MAXIMUM EXTERNAL LOAD. The maximum external load (including the weight of the cargo sling) for which authorization is requested may not exceed the rated capacity of the quick-release should not be attached in such a manner as to provide sling-leg angles of less than 30 degrees.

102. MINIMUM SLING-LEG CABLE STRENGHT. An analysis which considered the effects of 30-degree sling angles showed that the minimum cable strength design factor required would be 2.5 times the maximum external load for each leg regardless of the number of legs. Although this is the minimum strength required by Part 133, it may be desirable to double this value to allow for deterioration of the sling-leg cables in service.

This will result in a cable strength equal to 5 times the maximum external load.

Example: Maximum external load 850 pounds Minimum required sling-leg cable strength 850 × 2.5=2125 Minimum desired sling-leg strength 850 × 2.5 x 2=4250 A 3/16-inch, nonflexible 19-wire cable (MIL-W- 6940) provides a satisfactory cable strength. See figure 4.1, chapter 4, of AC 43.13-14 for a table of breaking strength of steel cable.

venience, the cable sizes desired for various loads have been calculated and are tabulated in figure

7.2 based on a factor of 5:

Aircraft Cable Size For Each Cargo Sling Leg Махітит MIL-C-5699 External Load and (pounds) MIL-W-6940 MIL-W-1511 MIL-C-5424 1/16 3/32 3/32

100. STATIC TEST. The cargo sling installation 200 3/32 1/8 1/8 must be able to withstand the static load required by FAR 133.43(a). Conduct the test as outlined 300 400 1/8 7/64 1/8 1/8 1/8 5/32 in Chapter 1 of this advisory circular. If re- 500 600 5/32 5/32 5/32 3/16 3/16 3/16 quired during the test, supports may be placed 700 3/16 3/16 3/16 at the landing gear to airframe attach fittings to 800 3/16 3/16 7/32 prevent detrimental deformation of the landing gear due to the weight of the aircraft. 1,000 1,200 900 3/16 7/32 7/32 7/32 7/32 1/4 7/32 7/32 1/4 1,400 1/4 1/4 9/32

101. SLING-LEG ANGLES OF CABLE-SUPPORTED 1,600 1/4 9/32 5/16 SUINGS. The optimum sling-leg angle (measured 1,800 &/16 5/16 5/16 from the horizontal) is 45 to 60 degrees. Mini- 2,000 5/16 11/32 3/8 mum tension in a sling leg occurs with a sling-leg angle of 90 degrees, and the tension approaches FIGURE 7.2.-Cable Load Table.

infinity as the angle approaches zero. 103. SLING INSTALLATION. Attach the cargo sling larger sling-leg angles are desirable from a stand- to landing gear members or other structure cappoint of cable strength requirements. Slings able of supporting the loads to be carried. Install 50 Chap 7

Rev. 1977 the quick-release device in a level attitude with the throat opening facing the direction as indicated on the quick-release device. When cables are used to support the quick-release device, make sure the cables are not twisted or allowed to twist in the direction to unlay the cable.

Some cargo release devices are provided with a fitting to permit installation of a guideline to assist in fully automatic engagement of the load target ring or load bridle. Secure the guideline to the quick-release device with a shear pin of a definite known value which will shear if a load becomes entangled on or over the guideline. Provision should also be made for cable-supported slings to be drawn up against the fuselage into a stowage position to prevent striking or dragging the release on the ground when not in use.

AC 43.13-2A

104. INSTALLATION OF RELEASE CONTROLS. See figure 7.3 for typical wiring diagram of the electrical controls.

a. Install a cargo release master switch, readily accessible to the pilot, to provide a means of deactivating the release circuit. The power for the electrical release circuit should originate at the primary bus. The "auto" position of the release master switch on some cargo release units provides for automatic release when the load contacts the ground and the load on the release is reduced to a preset value.

b. Install the cargo release switch on one of the pilot's primary controls. It is usually installed on the cyclic stick to allow the pilot to release the load with minimum distraction after maneuvering the load into the release position.

RELEASE SWITCH (ON CYCLIC CONTROL)

RELEASE MASTER SWITCH Automatic OFT OPEN

RELEASE

INDICATOR LIGHT

CIRCUTT

BREAKER

QUICK RELEASE DEVICE

PRIMARY BUS

AUTO-RELEASE

RELEASE MICROSWITCH SOLENOID (Closes when load 18 reduced to preset value - usually 80-120 1b.)

SAFETY

MICROSWITCH (Release closed) FIGURE 7.3.-Typical cargo sling wiring diagram.

Chap 7 51

AC 43.13-2A

c. Install the emergency manual release control in a suitable position that is readily accessible to the pilot or other crewmember. Allow sufficient slack in the control cable to permit complete cargo movement without tripping the cargo release.

d. The manual ground release handle, a feature of some cargo release units, permits opening of the cargo release by ground personnel.

e. Label or placard all release controls as to function and operation.

105. FUNCTIONAL TEST. Test the release action of each release control of the quick-release device with various loads up to and including the maximum external load. This may be done in a test fixture or while installed on the rotorcraft, if the necessary load can be applied.

If the quick-release device incorporates an automatic release, the unit should not release the load when the master switch is placed in the iautomatic" position until the load on the device is reduced to the preset value, usually 80 to 120 pounds.

FIGURE 7.4.--Typical cargo sling installation No. 1.

Rev. 1977

106. SUPPLEMENTAL FLIGHT INFORMATION. The aircraft may not be used in Part 133 externalload operations until a Rotorcraft-Load Combination Flight Manual is prepared in accordance with section 133.47 of that Part.

FIGURE 7.5.- -Typical cargo sling installation No. 1 (showing fuselage attachment fitting).

FIGURE 7.6-Typical cargo sling installation No. 1 (showing fore and aft limiting stops).

FIGURE 7.7.-Typical cargo sling installation No. 2 (cargo hook attached directly to underside of fuselage).

Chap 7

Rev. 1977 AC 43.13-2A FIGURE 7.8.-Typical cargo sling installation No. 3 (4-leg, cable suspended).

20.

FIGURE 7.9.-Typical cargo sling installation No. 3 (showing cable sling leg attachment to landing gear crosstube fitting).

FIGURE 7.10.- Typical cargo sling Installation No. 8 (showing cargo sling in stowed position).

107.-110. [RESERVED] Chap 7

AC 43.13-2A Rev. 1977 Section 2. CARGO RACKS

111. GENERAL. This section contains structural and design information for the fabrication and installation of a cargo rack used as an externalload attaching means for a Class A rotorcraftload combination operation under FAR Part 133.

112. FABRICATION OF CARGO RACKS. The type of construction and method of attachment depend upon the material to be used and the configuration of the rotorcraft involved. Illustrations of typical construction and installation methods are shown in figures 7.11-7.15.

113. STATIC TEST. The cargo rack installation must be able to withstand the static test load required by FAR 133.43(a). Conduct the test as outlined in chapter 1 of this handbook.

114. SUPPLEMENTAL FLIGHT INFORMATION. The aircraft may not be used in Part 133 externalload operations until a rotorcraft-load combination flight manual is prepared in accordance with section 133.47 of that Part.

FIGURE 7.12.-Typical cargo rack installation No. 1 (showing attachment detail).

FIoURE 7.11.-Typical cargo rack installation No. 1.

54 Chap 7

Rev. 1977 AC 43.13-2A FIGURE 7.13.-Typical cargo rack installation No. 2. FIGURE 7.15.-Typical cargo rack installation No. 3.

FIGURE_7.14.-Typical cargo rack installation No. 2 (showing rack partially installed).

115.-120. [RESERVED] Chap 7

6/1/79 AC 43.13-2A, CHG. 1 Chapter 8. GLIDER AND BANNER TOW-HITCH INSTALLATIONS Section 1. TOWPLANE CONSIDERATIONS [Deleted] Change 1 Chap 8 57

AC 43.13-2A, CHG 1 6/1/79 Section 2. TOW-HITCH INSTALLATIONS

126. STRUCTURAL REQUIREMENTS. The struc- * a. Glider tow hitches. Protection for the tural integrity of a tow-hitch installation on an towplane is provided by requiring use of a towline aircraft is dependent upon its intended usage. assembly which will break prior to structural Hitches which meet the glider tow criteria of this damage occurring to the towplane. The normal chapter are acceptable for banner tow usage. tow load of a glider rarely exceeds 80 percent of However, because the direction and magnitude of the weight of the glider. Therefore, the towline maximum dynamic banner towline loads occur assembly design load for a 1,000-pound glider within a more limited rearward cone of displace- could be estimated at 800 pounds. By multiplying ment than do glider towline loads, hitches which the estimated design load by 1.5 (to provide a safety meet the banner tow criteria of this chapter may margin), we arrive at a limit load value of 1,200 not be satisfactory for glider towing. Due to the pounds. The 1,200-pound limit load value is used in basic aerodynamic difference between the two static testing or analysis procedures per paragraph objects being towed, glider and banner tow-hitch 127 of this handbook to prove the strength of installations are treated separately with regard to the tow hook installation. When the hook and loading angles. structure have been proven to withstand* 20° Cone of Displacement FIGURE 8.1.-Glider tow angle.

SAFETY PAYS 10° Cone of Displacement * FIGURE 8.2. - Banner tow angle.

Chap 8

6/1/79 *the limit load, then the maximum breaking strength of the towline assembly is established at the design load of 800 pounds. Thus, the towline will break well before structural damage will occur to the towplane.

Another approach can be applied if the limit load carrying capabilities of a tow hook and fuselage are known. In this case, the known load value can be divided by 1.5 to arrive at the design load capabilities if the tow hook and fuselage limit loads are known to be 1,800 pounds. By dividing by 1.5 (1800 ÷ 1.5 = 1,200) we arrive at a design load value of 1,200 pounds. Thus, the maximum breaking strength of the towline assembly is established at 1,200 pounds and provides protection for the towplane.

Thus, in considering tow hook installations, one may establish maximum towline breaking strength

(1) Dividing the known limit load capabilities of the fuselage and two hook installation by 1.5; or

(2) Knowing the design load needs of the towline assembly and multiplying by 1.5 to arrive at a limit load. Then by analysis or static testing, determine that the hook and fuselage are capable of withstanding that limit load.

support a limit load equal to at least two times the operating weight of the banner.

127. STRUCTURAL TESTING. Adequacy of the aircraft structure to withstand the required loads can be determined by either static test or structural analysis.

a. Static testing.When using static tests to verify structural strength, subject the tow hitch to the limit load (per paragraph 126 a or b) in a rearward direction within the appropriate cone of displacement per figure 8.2. Testing to be done in accordance with the procedures of Chapter 1, paragraph 3, of this handbook.

b. Structural analysis. If the local fuselage structure is not substantiated by static test for the proposed tow load, using a method that experience has shown to be reliable, subject the fuselage to engineering analysis to determine that the local structure is adequate. Use fitting factor of 1.15 or greater in the loads for this analysis.

128. ATTACHMENT POINTS. Tow-hitch mechanisms are characteristically attached to, or at, tiedown points or tailwheel brackets on the air- Chap 8 AC 43.13-2A, CHG 1 frame where the inherent load-bearing qualities can be adapted to towing loads. Keep the length of the hitch-assembly arm from the airframe attachment point to the tow hook to a minimum as the loads on the attachment bolts are multiplied by increases in the moment arm.

129. ANGLES OF TOW. Tests should be conducted on the system at various tow angles to insure that:

a. There is no interference with the tailwheel or adjacent structure.

b. The towline clears all fixed and movable surfaces at the maximum cone of displacement and full surface travel.

c. The mechanism does not significantly decrease the clearance from the tailwheel to the

d. The tow hitch does not swivel. Experience has shown swiveling could result in fouling both the release line and towline during operations by the towplane.

e. The opened jaw of the hitch does not strike any portion of the aircraft.

130. PLACARDS. A placard should be installed in a conspicuous place in the cockpit to notify the pilot of the structural design limits of the tow system.

The following are examples of placards to be inbreaking strength not to exceed a. For glider tow-"Glider towline assembly pounds."

b. For banner tow-Tow hitch limited to maximum weight of pounds."

* Value established per paragraph 126 a (1) pr (2).

** Banner hitch limitations are one-half the load applied per paragraph 126 b.

131. WEIGHT AND BALANCE. In most cases, the weight of the tow-hitch assembly will affect the fully loaded aft c.g. location. To assure that the possibility of an adverse effect caused by the installation has not been ignored, enter all pertinent computations in the aircraft weight & balance records. (In accordance with the provisions contained in FAR 43.5(a) (4).)

132. TOW RELEASE MECHANISM.

a. Release lever. A placard indicating the direction of operation should be installed to allay the possibility of confusion or inadvertant opera-* 59

AC 43.13-2A, CHG 1 6/1/79 * tion, and the design of the release lever should provide the following:

(1) Convenience in operation.

(2) Smooth and positive release operation.

(3) Positioned so as to permit the pilot to exert a straight pull on the release handle.

(4) Sufficient handle travel to allow for normal slack and stretch of the release cable.

(5) A sufficient handle/lever ration to assure adequate release force when the towline is under high loads. (See fig. 8.3)

(6) Protection of cables from hazards such as:

(a) Wear and abrasion during normal operation.

(b) Binding where cables pass through fairleads, pulleys, etc.

(c) Accidental release.

(d) Interference by other aircraft com-

(e) Freezing and moisture accumulation when fixed or flexible tubing guides are used.

b. Test of the release. A test of the release and hook for proper operation through all angles of critical loading should be made using the design load for the glider or banner.

c. Release cable. Representative size strength characteristics of steel release cable are as shown in figure 8.4; however, it is recommended that all internally installed release cables be ½6-inch or larger.

Nonflexible Flexible Carbon Steel Carbon Steel 7 x 7 and 7 x 19 1 x 7 and 1 x 19 (MIL-W-1511A and Diameter (MIL-W-6904B) MIL-C-5424A) inches Breaking strength (Los.) Pounds 100 f. Breaking strength (lbs.) Pounds 100 p.

1/32 185 .20 3/64 375 .55 1/16 500 .85 480 .75 5/64 800 1.40 - 3/32 1,200 2.00 920 1.60 FIGURE 8.3-Typical tow-hitch release handle. FIGURE 8.4.- Representative steel cable qualities.

133. [Deleted] Change 1.

FIGURE 8.5.- [Deleted] Change 1. * 60 Chap 8 U.S. GOVERNMENT PRINTING OFFICE: 1979 O--302-515

(DELETED) -- Change 2 Chap. 8 61