CHAPTER 7. ROTORCRAFT EXTERNAL-LOAD-DEVICE
INSTALLATIONS CARGO SLINGS AND EXTERNAL RACKS SECTION 1. GENERAL
700. PURPOSE. This section contains structural and design information for the fabrication and installation of a cargo sling used as an external load attaching means for a Class B rotorcraft-load combination operation under Title 14 of the Code of Federal Regulations (14 CFR) part 133. As an external-load attaching means, a "cargo sling" includes a quick-release device (hook) and the associated cables, fittings, etc., used for the attachment of the cargo sling to the rotorcraft.
Part 133, § 133.43(d) specifies the requirements for the quick-release device.
701. HAZARDS AND WARNINGS. Particular attention should be paid to the effect of the sling load on the lateral as well as the fore and aft center of gravity (CG) of the rotorcraft. Use TC PMA, T50 or Designated Engineering Representative approved parts.
702. REFERENCES. It is the installer’s responsibility to ensure that the latest revisions of any regulations, advisory circulars, manufacturer’s data, etc., are used in any cargo sling or rack installation. Reference Title 14 of the Code of Federal Regulations part 27 and Civil Aviation Regulation 6, Rotorcraft Airworthiness; Normal Category.
703. INSTALLATION CONSIDERATIONS.
a. Location of Cargo Release in Relation to the Rotorcraft's CG Limits. An ideal location would be one that allows the line of action to pass through the rotorcraft's CG at all times. (See Figure 7-1, Illustration A.) However, with 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 CG (See Figure 7-1, Illustration B.)
(1) Whenever the line of action does not pass through the rotorcraft's CG due to the attachment method used, acceleration forces, or aerodynamic forces, the rotorcraft-load combined CG will shift from the rotorcraft's CG position.
Depending upon the factors involved, the shift may occur along either or both the longitudinal or lateral axes. The amount of shif t 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 CG (See Figure 7-1, Illustration C.) If the rotorcraft-load CG is allowed to shift beyond the rotorcraft's approved CG limits, the rotorcraft may become violently uncontrollable.
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FIGURE 7-1. LOCATION OF CARGO RELEASE IN RELATION TO THE ROTORCRAFT’S
CENTER OF GRAVITY
(2) Thus, any attachment method or location which will decrease the length of the arm will reduce the distance that the combined CG will shift for a given load (F) and line of action angle.
(See Figure 7-1, Illustration D.)
b. 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 device.
704. FABRICATION AND INSTALLATION.
a. Static Test. The cargo sling installation must be able to withstand the static load required by § 133.43(a). Conduct the test as outlined in chapter 1. If required during the test, supports may be placed at the landing gear to airframe attach fittings to prevent detrimental deformation of the landing gear due to the weight of the aircraft.
b. Sling-Leg Angles of Cable-Supported Slings. The optimum sling-leg angle (measured from the horizontal) is 45 to 60 degrees. Minimum tension in a sling leg occurs with a sling-leg angle of 90 degrees, and the tension approaches infinity as the angle approaches zero. Thus, larger sling-leg angles are desirable from a standpoint of cable strength requirements. Slings should not be attached in such a manner as to provide sling-leg angles of less than 30 degrees.
c. Minimum Sling-Leg Cable Strength.
(1) 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 Par 703 Page 62
will result in a cable strength equal to five times the maximum external load.
Example: Maximum external load 850 pounds Minimum required sling-leg cable strength 850 x 2.5 = 2125 Minimum desired sling-leg cable strength 850 x 2.5 x 2 = 4250
(2) A 3/16-inch, nonflexible 19-wire cable (MIL-W-6940) provides a satisfactory cable strength. For convenience, the cable sizes desired for various loads have been calculated and are tabulated in Table 7-1 based on a factor of 5:
TABLE 7-1. CABLE LOAD TABLE Aircraft Cable Size For Each Cargo Sling Leg
Maximum External Load (pounds)
MIL-C-5693 and MIL-W-6940
MIL-W-1511
MIL-C-5424 100 200 300 400 500 600 700 800 900 1,000 1,200 1,400 1,600 1,800 2,000 1/16 3/32 7/64 1/8 5/32 5/32 3/16 3/16 3/16 7/32 7/32 ¼ 1/4 5/16 5/16 3/32 1/8 1/8 1/8 5/32 3/16 3/16 3/16 7/32 7/32 1/4 1/4 9/32 5/16 11/32 3/32 1/8 1/8 5/32 3/16 3/16 3/16 7/32 7/32 7/32 ¼ 9/32 5/16 5/16 3/8
d. Sling Installation.
(1) Attach the cargo sling to landing gear members or other structure capable of supporting the loads to be carried. Install 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 un-lay the cable.
(2) 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.
e. Installation of Release Controls. See Figure 7-2 for typical wiring diagram of the electrical controls.
(1) 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.
(2) 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.
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FIGURE 7-2. TYPICAL CARGO SLING WIRING DIAGRAM
(3) 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.
(4) The manual ground release handle, a feature of some cargo release units, permits opening of the cargo release by ground personnel.
(5) Label or placard all release controls as to each function and operation.
f. Functional Test.
(1) 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.
(2) If the quick-release device incorporates an automatic release, the unit should not release the load when the master switch is placed in the "automatic" position until the load on the device is reduced to the preset value, usually 80 to 120 pounds.
FIGURE 7-3. TYPICAL CARGO SLING INSTALLATION NO. 1
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g. Supplemental Flight Information. The aircraft may not be used in part 133 external-load operations until a Rotorcraft-Load Combination Flight Manual is prepared in accordance with § 133.47 of that part. Appropriate entries should also be made in the aircraft’s weight and balance data, equipment list, and logbooks. The FAA Form 337, Major Repair and Alteration (Airframe, Powerplant, Propeller, or Appliance), should also be executed as required by part 43, § 43.5(a) and (b).
h. Inspection and Maintenance. Inspection of the complete installation should be accomplished prior to each lift for security and functionality.
Maintenance should be accomplished in accordance with the manufacturers instructions.
705. THRU 706. RESERVED
FIGURE 7-4. TYPICAL CARGO SLING INSTALLATION NO. 1 (SHOWING FUSELAGE ATTACHMENT FITTING)
FIGURE 7-5. TYPICAL CARGO SLING INSTALLATION NO. 1 (SHOWING FORE AFT LIMITING STOPS)
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FIGURE 7-6. TYPICAL CARGO SLING INSTALLATION NO. 2 (CARGO HOOK ATTACHED
DIRECTLY TO UNDERSIDE OF FUSELAGE)
FIGURE 7-7. TYPICAL CARGO SLING INSTALLATION NO. 3 (4-LEG, CABLE SUSPENDED)
FIGURE 7-8. TYPICAL CARGO SLING INSTALLATION NO. 3 (SHOWING CABLE SLING
LEG ATTACHMENT TO LANDING GEAR CROSSTUBE FITTING)
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Par 704 Page 67 (and 68) FIGURE 7-9. TYPICAL CARGO SLING INSTALLATION NO. 3 (SHOWING CARGO SLING
IN STOWED POSITION)
Par 707 Page 69 SECTION 2. CARGO RACKS
707. GENERAL. This section contains structural and design information for the fabrication and installation of a cargo rack used as an external-load attaching means for a Class A rotorcraft-load combination operation under part 133.
708. FABRICATION OF CARGO RACKS. The type of construction and method of attachment depends upon the material to be used and the configuration of the rotorcraft involved. Illustrations of typical construction and installation methods are shown in Figures 7-10 through 7-14.
FIGURE 7-10. TYPICAL CARGO RACK INSTALLATION NO. 1
709. STATIC TEST. The cargo rack installation must be able to withstand the static test load required by § 133.43(a). Conduct the test as outlined in chapter 1.
710. SUPPLEMENTAL FLIGHT INFORM- ATION. The aircraft may not be used in part 133 external-load operations until a rotorcraft-load combination flight manual is prepared in accordance with § 133.47.
711. THRU 799. RESERVED
FIGURE 7-11. TYPICAL CARGO RACK INSTALLATION NO. 1 (SHOWING ATTACHMENT DETAIL)
FIGURE 7-12. TYPICAL CARGO RACK INSTALLATION NO. 2
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Par 710 Page 71 (and 72) FIGURE 7-13. TYPICAL CARGO RACK INSTALLATION NO. 2 (SHOWING RACK PARTIALLY INSTALLED)
FIGURE 7-14. TYPICAL CARGO RACK INSTALLATION NO. 3
CHAPTER 8. GLIDER AND BANNER TOW-HITCH
INSTALLATIONS
800. PURPOSE. This chapter contains design and installation information for banner and glider tow hitches. Guidance for inspection, service, and continuous airworthiness requirements for the hitches are also addressed in this chapter.
801. HAZARDS AND WARNING. The direction and maximum arc of displacement of banner towline loads occur within a more limited rearward cone of displacement than do g lider towline loads (see Figure 8-1 and 8-2). Hitches that meet the banner tow criteria of this chapter may not be suitable for glider towing. Due to the basic aerodynamic differences between the two objects being towed, glider and banner tow-hitch installations are treated separately with regard to loading angles.
802. REFERENCES. This chapter will occasionally reference manufacturer’s documents, certain parts of Title 14 of the Code of Federal Regulations (14 CFR), and certain FAA advisory circulars (AC). It is the installer’s responsibility to ensure the latest revisions of these documents are used as reference material. Refer to Civil Aviation Regulation 6, Rotorcraft Airworthiness; Normal Category, and Order 8900.1, Volume 8, Chapter 5, Section 2, Changes to Special Purpose for Restricted Aircraft.
803. INSTALLATION CONSIDERATIONS.
a. Weight and Balance. In most cases, the weight of the tow-hitch assembly will affect the aft center of gravity (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 and balance records, in accordance with the provisions contained in 14 CFR part 43, § 43.5(c). The requirements of § 43.5(a) and (b) should also be addressed, for maintenance record entries and repair and alteration forms.
b. Equipment List. The aircraft equipment list should be updated to reflect any tow equipment installations. Consideration should also be given to adding appropriate revisions to the aircraft’s Pilot Operating Handbook (POH), or Flight Manual, as required.
c. Corrosion Protection. Tow hitches are traditionally simple mechanical devices; however, improper care may lead to hazardous conditions for the aircraft operator. The aft, external location of tow hitches exposes them to the elements, and proper corrosion protection methods should be employed to prevent improper operation or malfunction.
Reference should be made to the manufacturers documents or current edition of AC 43.13-1, Acceptable Method, techniques, and Practices―Aircraft Inspection and Repair, chapter 6, for some additional corrosion protective measures that may be used.
804. FABRICATION AND INSTALLATION PROCEDURES.
a. Methods. Accomplish the installation of tow hitches using manufacturer’s data or that data previously approved by a representative of the Administrator, when available, such as a Supplemental Type Certificate (STC) or a Field Approval. Installations requiring fabrication of brackets, parts, fittings, etc. should be accomplished using data in the form of ACs such as AC 43.13-1 (current edition), to determine material requirements, load requirements, and type and size of hardware used.
b. Structural Requirements. The structural integrity of a tow-hitch installation on an aircraft is dependent upon its intended usage.
c. Attachment Points . Tow-hitch mechanisms are characteristically attached to, or at, Par 800 Page 73
tie down points or tail wheel brackets on the airframe. These are points where the design loadbearing qualities may be sufficient for towing loads.
Keep the length from the airframe attachment point to the tow hook at a minimum as the loads on the attachment bolts are multiplied by increases in the attachment arm.
d. Glider Tow Hitch Load Testing.
(1) Protection for the tow-plane is provided by requiring use of a towline assembly that will break prior to structural damage occurring to the tow plane. The normal to w load imposed on the hitch rarely exceeds 80 percent of the weight of the glider. Therefore, the towline assembly design load for a 1,000-pound glider could be estimated at 800 pounds. By multiplying the estimated design load by 1.5 (to provide a safety margin), we arrive at a limit load value of 1,200 pounds. The 1,200-pound limit load value is used in static testing or analysis procedures per paragraph 8-2 to prove the strength of the tow hook installation. When the hook and its attachment to the aircraft structure have been proven to withstand the limit load, 1,200 pounds in this example, 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.
FIGURE 8-1. GLIDER TOW ANGLE
FIGURE 8-2. BANNER TOW ANGLE
(2) 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 limits. For example, if the tow hook and fuselage limit loads are known to be 1,800 pounds, divide this by 1.5 (1,800 ÷ 1.5 = 1,200) and we arrive at a design load value of 1,200 pounds. Thus, if a towline assembly rated at 1,200 pounds is used it will break before reaching the structural limits of the hook and attaching structure.
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