10-13 Figure 10-7. Ceiling Mounted Inertia Reel―Double Over-the-Shoulder Type Harness Fig ure 10-8. The Four-Point Independent Harnesses Positioning
10-14 Figure 10-9. The Five-Point Harness Positioning Fig ure 10-10. “Y” Type Retractor Positioning
10-15
10.15 Area and Angle Deviations. While the areas and angles given in the above paragraphs are intended to assist in the selection of attachment points, they should be considered the desirable optimum. Area or angle deviations, or both, could decrease the overall efficiency of the restraint system. However, they may be necessary in order to permit harness installation in an aircraft that cannot accommodate ideal harness geometry. As discussed in Section 3, Static Strength and Testing, angles have a significant bearing on static test load requirements.
10.16 Structural Attachment. This chapter presents only a few generic examples of structural attachment of harnesses that may be determined to be acceptable for a specific application. See Section 3 for strength requirements of the attachment.
10.16.1 Floor and Seat Attachments. The dual over-the-shoulder or “Y” type harness may be used with either floor or seat mounting points. Typical installations are illustrated in Figures 10-11, Floor Mounted Inertia Reel―Double Over-the-Shoulder Type Harness;
10-12, Typical Floor Mounted Inertia Reel Installation; and 10-13, Typical Seat Mounted Inertia Reel Installation. Several factors need to be considered that may make this configuration undesirable.
10.16.1.1 The floor, seat structure, and anchorages must be capable of withstanding the additional “G” loads imposed upon them by the restraints.
10.16.1.2 The height of the seat back should at least be equal to the shoulder height of the seated occupant to reduce the possibility of spinal compression injuries.
10.16.1.3 Harness guides may be necessary to maintain proper harness position over the occupant’s shoulders.
10.16.1.4 Seat back strength is critical to performance of these installations. Folding seat backs must have a locking mechanism that can withstand the imposed loads;
however, a locked seat back may impede emergency egress of other occupants and create an impact hazard for occupants seated aft of the locked seat back.
10-16 Figure 10-11. Floor Mounted Inertia Reel―Double Over-the-Shoulder Type Harness Fig ure 10-12. Typical Floor Mounted Inertia Reel Installation
10-17 Figure 10-13. Typical Seat Mounted Inertia Reel Installation
10.16.2 Airframe Attachments. The method used for the attachment of shoulder harness anchorages is dependent upon the construction features of the aircraft involved. When selecting a point of attachment, careful consideration should be given to the static strength and testing requirements found in Section 3.
10.16.2.1 Monocoque/Semimonocoque Type Constructions. Illustrations of typical aircraft members and installation methods are shown in Figures 10-14, Typical Wing Carry-Through Installation, through 10-18, Typical Localized Stringer Section Reinforcement Installation. Intercostal doublers and stiffeners are frequently added to provide increased strength. AC 43.13-1, Acceptable Methods, Techniques, and Practices—Aircraft Inspection and Repair, Chapter 4, Metal Structure, Welding, and Brazing, provides useful information on fabrication and installation of sheet metal repairs that may be applicable to doubler and stiffener installations.
10.16.2.2 Tube Type Construction. Various typical methods of attaching shoulder harness anchorages are shown in Figure 10-19, Typical Attachments to Tubular Members. Attachment should be accomplished at the intersection of tubular members and not at the center of single unsupported tubes. When aircraft cable is used as a component in a shoulder harness anchorage, swage the cable terminals in accordance with the procedures contained in AC 43.13-1, Chapter 7, Aircraft Hardware, Control Cables, and Turnbuckles.
10-18
10.16.3 Structural Repair Instructions. In some instances, structural repair instructions are provided in the aircraft manufacturer’s maintenance manual. While these instructions are primarily intended for use in repairing defective or damaged structure, they may also be used as reinforcement methods for shoulder harness attach fittings.
10.16.4 Flexible Attachments. Various aircraft are designed so that fuselage members or skin, or both, will flex or “work.” This type of structure should not be heavily reinforced for the purpose of attaching shoulder harnesses, as this would defeat the design purpose. In these cases, use a localized reinforcement, such as that shown in Figure 10-18, at the attachment point. This will allow the fuselage to flex while still maintaining a collapsible structure to absorb the loads encountered in a crash.
Figure 10-14. Typical Wing Carry-Through Installation Fig ure 10-15. Typical Hat Section Reinforcement Installation
10-19 Figure 10-16. Typical Bulkhead Reinforcement Installation Fig ure 10-17. Typical Stringer Section Reinforcement Installation
10-20 Figure 10-18. Typical Localized Stringer Section Reinforcement Installation Fig ure 10-19. Typical Attachments to Tubular Members
10-21 Section 3. Static Strength and Testing
10.17 Proof of Compliance. Retrofit harness installations must be proven to meet minimum strength requirements applicable to the aircraft certification basis; otherwise, their ability to provide adequate restraint under accident decelerations is undetermined. Static strength requirements can be demonstrated by static testing or stress analysis.
10.17.1 Static tests are accomplished by application of the ultimate load to the harness attachment point. Since TSO and PMA restraints have already been shown to comply with strength requirements, static tests are accomplished to substantiate the airframe attachment. Static testing in operational aircraft is undesirable because there is the risk of airframe damage;
however, static testing in a conforming but unserviceable fuselage or cabin section of the same M/M aircraft is an acceptable alternative.
10.17.2 Stress analysis is accomplished through engineering data developed for the specific attachment point. This data can be developed by FAA DERs with aircraft structural approval authority. The DER will perform the analysis, prepare a report, and complete FAA Form 8110-3, which demonstrates that a specific regulation has been met. The DER will list on FAA Form 8110-3 the data needed to show the structural adequacy of the installation of the harness. Since the DER is only approving the structural aspects of harness attachment, approval for return to service for the installation as a whole is accomplished by the authorized individual or agency on FAA Form 337. To locate an appropriately authorized DER in your locality, contact the FAA Flight Standards District Office (FSDO) or the responsible Aircraft Certification Service office for the area.
Note: It is up to the installer to determine which method will be the most efficient. For simple installations, it is possible that the DER could perform stress analysis in just a few hours. The DER may also be able to assist in the design of attachment points.
10.18 Load and Distribution―Airplanes. Table 10-1 provides a reference to determine minimum static test loads based upon certification date, category, standard occupant weight, fitting factors, and harness factors. The numbers in each line are multiplied together to obtain the static test load minimum standard.
10.18.1 Load. For airplanes type certificated prior to September 14, 1988, occupant weights were established as 170 pounds for normal category and 190 pounds for utility and acrobatic category airplanes. The additional 20 pounds accounts for a parachute. For aircraft type certificated on or after this date, 215 pounds is the established weight for the three categories.
10.18.2 Distribution. In the assessment of a combined shoulder harness-safety belt restraint system, a forward static test load distribution of 40 percent to the shoulder harness and 60 percent to the safety belt has been an acceptable combined static test load distribution.
Therefore, demonstration of harness attachment static strength is accomplished by applying 40 percent of the ultimate restraint system load to a single or “Y” type harness attachment or 20 percent of the load simultaneously to each attachment of a dual
10-22 independent harness. Static strength test load requirements for the fifth belt of a five-point system have not been established for aircraft.
10.18.3 Factor. The airplane must be designed to protect each occupant during emergency landing conditions. The critical ultimate static load is established at 9.0 Gs in the forward direction. If the attachment can sustain this load, it is assumed the lesser loads in upward, sideward, and downward directions will be accommodated.
10.18.4 Fitting Factors. Regulations prescribe multiplying factors of safety, fitting factors, that structural attachments must be subjected to in order to prove minimum load bearing capabilities. A 1.15 fitting factor for restraint attachments in small aircraft type certificated prior to September 14, 1969, must be used. For aircraft type certificated on or after this date, a fitting factor of 1.33 must be used.
10.19 Load and Distribution―Normal Category Helicopters.
10.19.1 Load. For helicopters type certificated in the normal category, occupant weights are established as 170 pounds.
10.19.2 Distribution. In the assessment of a combined shoulder harness-safety belt restraint system, for helicopters type certificated prior to December 13, 1989, a forward static test load distribution of 40 percent to the shoulder harness and 60 percent to the safety belt is the required combined static test load distribution.
10.19.3 G Factor. For helicopters type certificated prior to December 13, 1989, the critical ultimate static load is established at 4.0 Gs in the forward direction. If the attachment can sustain this load, it is assumed the loads in upward, sideward, and downward directions will be accommodated.
Note: Due to increased G factors and limited applicability for retrofit harness installations, helicopters type certificated on or after December 13, 1989, are not addressed in this chapter. Additionally, airplanes seat/restraint systems type certificated after September 14, 1988, must undergo dynamic testing, as prescribed by 14 CFR § 23.562, which is not covered in this chapter.
10.20 Demonstrating Static Loads. Two methods are available to demonstrate compliance with static load requirements. One method involves the fabrication of test blocks and application of the ultimate load to the blocks that are being restrained by the harness. The other method involves use of an equation to calculate and simulate the ultimate load at the harness attachment fitting, compensating for horizontal and vertical angles as a result of restraint attachment geometry.
10.20.1 Test Blocks. Although AC 23-4, Static Strength Substantiation of Attachment Points for Occupant Restraint System Installations, was cancelled on January 27, 2003, it is still available in various archives and can be used as an acceptable method for fabricating test blocks and applying static loads to restraints. Using this method, deceleration loads imposed upon the restraint by the occupant are accurately duplicated, and the load figures in Table 10-1 can be applied directly to the test blocks. Because of the time-consuming
10-23 nature of fabricating test blocks for this procedure, it may be desirable to use the second method described below that, in effect, replaces the test blocks with an equation.
10.20.2 Calculation. The alternate method is to perform a mathematical calculation to account for the angles that the attachment points are subjected to during decelerations. Increases in angles can result in a significant increase in the load that must be applied to simulate belt loading during occupant decelerations. It is not uncommon for a harness to impose both vertical and horizontal angles that must be compensated for by use of the equation (see Figures 10-20, Vertical Angle, and 10-21, Horizontal Angle). In the case of a “Y” type harness, the horizontal angle is eliminated if the attachment is located along the vertical centerline of the occupant.
10.20.2.1 The equation used to perform the calculation is:
Test Load = [Static Test Minimum] [1/Cosine(vertical angle)] [1/Cosine(horizontal angle)]
10.20.2.2 Table 10-2, Calculated 1/Cosine Values, provides 1/Cosine numbers for various angles. These numbers can be inserted into the equation to determine the static load that needs to be applied. For example, using Tables 10-1 and 10-2, if the airplane is normal category, post-September 14, 1969, and pre-September 14, 1988, the test load is 814 pounds. If the horizontal angle is 10 degrees and the vertical angle is 15 degrees, the belt load is calculated as follows.
Test Load = [814][1/Cos(15)][1/Cos(10)] = [814][1.035][1.015] = 855 pounds This is not a large increase, but if both angles were 30 degrees, the belt load would be 1,086 pounds.
[814][1.155][1.155] = 1,086 If there is no angle present in one of the directions, that portion of the equation is eliminated. For an airplane type certificated in the normal category prior to September 14, 1969, with a vertical angle of 25 degrees, the equation would be: [704][1.103] = 777 pounds.
10.20.3 After obtaining the test load value, the load is applied for a minimum of 3 seconds forward (horizontally) when using test blocks or in the direction of the restraint angles when applying the load to the attachment point(s).
10.20.4 Equipment used for measuring test loads must be calibrated to a standard acceptable to the FAA. Load cells can perform these measurements; however, other equivalent equipment as determined by the user may be acceptable.
10-24 Table 10-1. Static Test Requirements TC Basis Airplanes Category G Factor Weight Factor (pounds) Fitting Factor Harness Factor (40%) Static Test Minimum Load (pounds) Effective 09/14/88 and on N, U, A 9 215 1.33 0.4 1,029 Effective 09/14/69 N 9 170 1.33 0.4 814 U, A 9 190 1.33 0.4 910 Pre 09/14/69 N 9 170 1.15 0.4 704 U, A 9 190 1.15 0.4 787 Helicopter Pre-12/16/84 N 4 170 1.15 0.6 469 Helicopter Post-12/16/84 Pre-12/13/89 N 4 170 1.33 0.6 543 Note 1: N = Normal, U = Utility, A = Acrobatic Note 2: Airplane seat/restraint systems type certificated after September 14, 1988, must undergo dynamic testing, as prescribed by 14 CFR § 23.562, which is not covered in this chapter.
Table 10-2. Calculated 1/Cosine Values Degrees 5 10 15 20 25 30 35 40 50 60 1/Cos Multiplier 1.004 1.015 1.035 1.064 1.103 1.155 1.221 1.305 1.556 2.0