CHAPTER 11. ADDING OR RELOCATING INSTRUMENTS 1100. PURPOSE. This chapter contains structural and design guidance to consider when aircraft alterations are to be accomplished by the addition or relocation of instruments.
1101. HAZARDS AND WARNINGS. The rapid advance of technical progress in the aviation industry has resulted in a virtual explosion of hightech aftermarket instrumentation. In many instances, these innovations are marketed without the support of a proper certification process. As a result, much of the equipment purported to be the most advanced may not be certificated or even certificable for installation on aircraft for which uninformed purchasers have chosen the equipment for installation.
NOTE: The burden of provision of and compliance with approved data falls entirely on the technician approving the aircraft for return to service.
1102. ADDITIONAL REFERENCES. Before initiating an alteration involving the addition or relocation of instruments, the regulatory basis of the aircraft must be considered. The introduction page of this document provides guidance to be considered on the subject of certification basis. For this chapter, the following additional references are provided:
a. Department of Commerce, Aeronautics Branch, Aeronautics Bulletin 7a, section 73(A) through (H) provides regulatory requirements for electrical equipment applicable to the 24 makes and 115 models of aircraft certificated under the provisions of that document. Section 75(B) of the same document provides additional regulatory requirements for instrumentation to be installed on the aircraft listed therein.
NOTE: Exercise care in the use of Bulletin 7a as certification basis proof data. Subsequent changes in ownership and production of some makes and models may have also resulted in changes of certification basis.
b. Department of Commerce, Civil Aeronautics Administration Bureau of Regulation Inspection Handbook, as revised to January 29, 1947, chapter XVII provides a listing of approved type certificated and 609 Group 2 aircraft under those categories. The Group 2 memoranda are identified and briefly described but complete text documentation is not provided. As of this publication no known repository of complete Group 2 memoranda has been identified.
c. Civil Aeronautics Regulation (CAR) 18 and the related Civil Aeronautics Manual (CAM) 18 established and maintained the standards and practices for all maintenance, repair, and alteration of aircraft applicable under the provisions of the CAR. That document remained in effect until the advent of the systemic changes dictated by the Federal Aviation Act of 1958 and the resultant documentary changes.
d. CAR 6, Rotorcraft Airworthiness; Normal Category.
1103. PREPARATION. Before initiating any alteration involving the addition or relocation of instruments the regulatory basis of the aircraft must be considered. For the purposes of this chapter the following references are provided:
• Civil Air Regulation (CAR) 3.661 through 3.676 provides the regulatory basis for installation of instruments in CAR 3 airplanes
• CAR 4a532 through 4a537 provide the regulatory basis for supplemental Par 1100 Page 123
variations for installation of instruments in CAR 4 non-air carrier (NAC) airplanes
• Title 14 of the Code of Federal Regulations (14 CFR) part 23, § 23.1321 provides the regulatory basis for the arrangement and visibility of instruments in part 23 airplanes
a. Structure. Chapter 1 provides guidance by which structural integrit y may be determined.
Chapter 2, paragraphs 202 and 203, provides information pertinent to instrument panel installations.
b. Location. In the absence of specific regulatory requirements, for installation of instruments required for operation under Instrument Meteorological Conditions (IMC), the recommended configuration is in the form of the basic “T” (Figure 11.1).
FIGURE 11-1. THE BASIC “T” VMC
INSTRUMENT CONFIGURATION
(1) The aircraft attitude indicator is located at the top center of the installation.
(2) The air speed indicator is located adjacent to and left of the attitude indicator.
(3) The altimeter is located adjacent to and to the right of the attitude indicator.
(4) The directional indicator is located adjacent to and immediately below the attitude indicator.
NOTE: As instrument panel installations have become more complex, the basic “T” has become more difficult to accomplish. With the relocation of the directional indicator, the relationship of the ASI-ATT-AI combination is assumed.
1104. INSTALLATION. Throughout the development of the complexity of aircraft the requirements for the mounting of all instruments has remained the same. A ll instruments must be mounted so they are visi ble to the crewmember primarily responsible for their use.
a. Structure. Before initiating changes to instrument panel installations, determine the load bearing requirements of the component. In some aircraft the instrument panel is stressed as the primary structure; with the majority, the panel is treated as the secondary structure with load bearing responsibilities limited to instruments and equipment installed. Regardless of the structural nature, the altering technician is responsibility for the installation to its original or properly altered state.
(1) In all cases, where available, refer to the manufacturer’s instructions for continued airworthiness (ICA) when considering alterations to any structure. For aircraft manufactured before the regulatory requirement for the manufacturer to provide such information, refer to AC 43.13-1, Acceptable Methods, Techniques, and Practices (as amended), chapter 4, section 4 and/or chapter 2, of this AC for methods and techniques of retaining structural integrity.
(2) Failing availability or applicability of those sources we recommend that the altering individual have a Designated Engineering Representative (DER) approve the data before accomplishing of the task.
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b. Instrument Plumbing. The majority of CAR 3 and CAR 4 general aviation aircraft use instrumentation identified as “wet” gauges. This term has nothing to do with what is being measured by the gauge, but rather the material to be measured is delivered directly to the gauge. Typical “wet” systems are fuel pressure, engine oil pressure and, when so equipped, hydraulic pressure. While not usually considered “wet” systems, the airspeed and vacuum systems use this same system. The tubing used is generally thin walled seamless aluminum or stainless steel. However, flexible rubber and/or neoprene lines are prevalent in airspeed and vacuum sensing lines.
(1) The most reliable source of information on the material used in a particular aircraft is the manufacturer’s ICA. If that source is not available, AC 43.13-1 (as amended), chapter 8 provides inspection and maintenance guidance on fuel lines, while chapter 9 provides the same level of information on hydraulic systems.
(2) The installation of remote sensing systems per part 23 has replaced the “wet” systems with system pressures being remotely measured by transducers and electrica lly transmitted to the cockpit indicators.
(3) Aircraft with long histories of alteration may contain a mixture of mechanical and/or electrical analogue instruments and digital presentations.
c. Venturi Vacuum Source. A venturi is a relatively low-cost means of producing the vacuum to operate gyroscopic instruments such as the turn and bank, directional gyro, and artificial horizon. It is mounted on the exterior of the fuselage, parallel with the longitudinal axis of the aircraft. As the aircraft moves through the air, an area of reduced pressure, a partial vacuum, is created in the throat of the venturi. The venturi throat is connected by a tube to the gyro instrument case, resulting in a reduction of pressure within the case. Ambient air, entering the instrument through the inlet port of the case, passes over the gyro rotor, causing it to spin rapidly and causes the instrument to activate.
Venturis are available in sizes providing 2”, 4”, and 9” Hg. of vacuum.
(1) The vacuum requirement for the operation of each instrument is generally provided on the data plate attached to its case. If it is not, obtain the data from the manufacturer or another acceptable source, such as AN and MS standards.
(2) If the vacuum available exceeds the requirements of the installation it will be necessary to install an in-line regulator to adjust the flow within requirements.
d. Pumped Vacuum Source. Current aviation technology has produced durable lightweight, mechanically-driven, pump and control systems for providing instrument operating vacuum.
Selection of the correct pump remains dependent on the vacuum requirements of the system being driven. Control is provided by either the system itself or an in-line regulator. Production and aftermarket standby vacuum systems are available to provide operating capabilities in the event of failure of the primary mechanically-driven system.
e. Calculating Vacuum Loads. When a venturi vacuum source is selected, do not assume the venturi selected will provide sufficient flow and negative pressure to operate the instrument package.
Even within the make and model of the venturi selected, tolerance may be insufficient to meet the requirements of the system selected. Therefore, it is essential that the vacuum load requirements be carefully evaluated.
(1) Gyroscopic instruments require optimum value of airflow to produce their rated rotor speed. For instance, a specific bank and pitch indicator required approximately 2.30 cubic feet per minute (CFM) flow and a resistance, or pressure drop, of 4” Hg. If the altering technician has selected a 2” Hg. Venturi, the resultant vacuum would be insufficient to drive the instrument.
(2) The following vacuum driven items are examples for the purposes of calculation:
TABLE 11-1. CFM/VACUUM VALUES instrument flow vacuum Bank and Pitch Indicator 2.3 cubic feet per minute 4.0” Hg Directional Gyro Indicator 1.3 cubic feet per minute 4.0” Hg Turn and Bank Indicator .5 cubic feet per minute 2.0” Hg Total flow required 4.10 cubic feet per minute The above listed instruments are listed in Tables 11-2 and 11-4. Optimum values are shown in Table 11-4. The negative pressure air source must deliver not only the optimum value of vacuum for the instruments, but must also have sufficient volume capacity to accommodate the total flow requirements of the various instruments it serves.
NOTE: The example components listed above and in Tables 11-2 and 11-4 are accepted as viable and the operating parameters related to them as accurate. The nomenclature and the properties related have no effect on the resulting computations.
TABLE 11-2. PROVIDES INSTRUMENT AIR CONSUMPTION FOR AN EXAMPLE VACUUM
DRIVEN AUTOPILOT AND INSTRUMENT INSTALLATION Air consumption at sea level Instrument Differential drop in. Hg suction (Optimum) Cubic feet/per minute Automatic Pilot System (Type A-2, A-3, & A-3A)
Directional gyro control unit across mount assembly 5.00 2.15* Bank & climb gyro control unit across mount assembly 5.00 3.85* Total ― 6.00* Automatic Pilot System (Type A-4) Directional gyro control unit 5.00 3.50* Bank & climb gyro control unit 5.00 5.00* Total ― 8.50* Bank & pitch indicator 4.00 2.30 Directional gyro indicator 4.00 1.30 Turn & bank indicator 2.00 .50 (*) NOTE.― Includes air required for operations of pneumatic relays.
TABLE 11-3. PROVIDES THE EQUIVALENT STRAIGHT TUBE LINE DROPS FOR THE 90-DEGREES ELBOWS INSTALLED IN ANY GIVEN SYSTEM (BY O.D.)
Tubing size Pressure drop in a 90˚ elbow in terms of length of straight tube equivalent to a 90˚ elbow O.D inch Wall thickness inch Feet 1/4 X 0.035 0.28 3/8 X 0.035 0.46 1/2 X 0.042 0.62 5/8 X 0.042 0.81 3/4 X 0.049 0.98 1 X 0.049 1.35
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Par 1104 Page 127 TABLE 11-4. PROVIDES THE DIFFERENTIAL PRESSURE ACROSS THE INSTRUMENT
INLET AND OUTLET OF THE EXAMPLE INSTRUMENT PACKAGE Suction in inches of Mercury (Hg) Instrument Minimum Optimum Maximum Automatic Pilot System (Type A-2, A-3, & A-3A) Directional gyro control unit across mount assembly 4.75 5.00 5.25 Bank & climb gyro control unit across mount assembly 4.75 5.00 5.25 Gauge reading (differential gauge in B & C control unit) 3.75 4.00 4.25 Automatic Pilot System (Type A-4) Directional gyro control unit 3.75 5.00 5.00 Bank & climb gyro control unit 3.75 5.00 5.00 Bank & pitch indicator 3.75 4.00 5.00 Directional gyro indicator 3.75 4.00 5.00 Turn & bank indicator 1.80 2.00 2.20
(3) Calculation of Equivalent Straight Line Pressure Drop Due to Routing. In addition to the effects of the instrumentation itself consideration must be given to the parasitic resistance to flow provided by the routing of the tubing providing the source pressure to the instruments. For this problem, assume the example system has four right-angle elbows and 20 feet of 1/ 2-inch O.D. X 0.042” tubing.
(4) Solution to the Stated Problem (Step-by-Step).
(a) Accept the flow requirements of the example instruments listed in paragraph 1104e(2) as 4.10 CFM. The pressure drop for one 90-degree 1/2-inch O.D. elbow is equivalent to 0.62 feet of straight tubing. Therefore, the pressure drop of the four 90-degree elbows is the equivalent to 2.48 feet of tubing. (Ref. Table 11-3.)
(b) Determine the pressure drop through 22.48 feet (20 feet plus the above 2.48 feet) of 1/2-inch O.D. X 0.42 inch tubing at 4.10 CFM flow. (See Figure 11-2.)
FIGURE 11-2. PRESSURE DROP DATA FOR SMOOTH TUBING
2.0 PRESSURE DROP PER TEN FOOT LENGTH - INCHES OF MERCURY *
1.0
o. TUBING SIZE 5/16
0.8 3/8 x .035 1/2 x .042 5/8 x .042 - 3/4 x .049
0.2 1 × „049
0.1
0.5 0.75 1 1.5 2 3 4 5 6 8 10 15 20
AIR FLOW - CUBIC FEET PER MINUTE
1. The pressure drop for each 10-foot length of that tube is 0.68" Hg.
2. Divide 22.48 feet of tubing to determine the number of 10-foot lengths (1.e.,
22.48 divided by 10 equals 2.248).
3. Multiply the number of sections (2.248) by 0.68" Hg to obtain the pressure drop through the system (0.68 X 2.248 = 1.53" Hg).
4. The pump must therefore be capable of producing a minimum pressure Par 1104
differential of 5.63” Hg (i.e., 4.10” Hg) for maximum instrument usage + 1.53” Hg (determined) at a flow of 4.10 CFM.
(c) Filters are used to prevent dust, lint, or other foreign matter from entering the instrument and vacuum system. Filters may be located at the instrument intake port or the manifold intake port when instruments are interconnected.
Determine if the flow capacity of the filter is equal to or greater than the flow capacity of the vacuum system. If it is not, the restriction will create a pressure drop in the system. AC 43.13-1 (as amended), chapter 12, paragraph 38a through d provides additional guidance on the installation and maintenance of venturi and engine-driven vacuum and instrument systems.
(d) Electrical supply for instruments. AC 43.13-1 (as amended), chapter 11 provides 16 sections of text on the inspection and installation of electrical systems.
(e) Instrument lighting. The regulatory requirements for instrument lighting have changed little as certifica tion specifications have developed.
1. CAR 3.696 and 3.697 provide two paragraphs simply stating that (1) instrument light installations should be safe, (2) provide sufficient illumination to make instruments and controls easily readable, and (3) shall be installed in such a manner as not to shine in the pilot’s eyes.
2. CAR 4a577 adds that instrument lighting shall be equipped with a rheostat control for dimming unless it can be shown a nondimming light is satisfactory.
3. Section 23.1381 consolidates the language of the earlier regulatory requirements with no appreciable changes.
(f) Magnetic Direction Indicator.
The magnetic compass is commonly referred to as the whiskey compass. The maximum deviation limitations of magnetic compasses have changed little since 1928.
1. CAR 3.666 specifies that the Magnetic Direction Indicator must be installed that its accuracy is not excessively affected by the airplane’s vibration or magnetic field. After compensating for this in the installation, the deviation level in flight must not exceed 10 degrees on any heading. CAR 3.758 provides and requires the installation and maintenance of the Compass Deviation Card.
2. CAR 4a562 refers to dampening and compensation of the installation and acknowledges the effects of electrical disturbances resulting from the increase in proliferation of aircraft electrical systems. As a result, specific aircraft required to meet the provisions of CAR 4a may be required to display two compass correction cards, one for operation with electrical power off, and one for electrical power on operation.
3. Section 23.1327 provides location and deviation considerations also provided for under CAR 3.666 and CAR 4a562. This section additionally provides for more than 10 degrees deviation of the non-stabilized magnetic compass due to the operation of adjacent high draw electrical components: if either a magnetic stabilized indicator, not having a deviation of greater than 10 degrees on any heading, or a gyroscopic direction indicator is installed.
(g) Alternative installations to liquid-filled Magnetic Direction Indicators. With the development of the vertical reading dry compass, it is practical to replace the liquid-filled compass with a viable and approved alternative.
1. Vertical reading dry compasses provide the pilot with a true relationship of the aircraft to azimuths as opposed to the reverse reading requirements of the liquid filled compass.
2. The vertical reading dry compass does not leak and requires no periodic refilling or cleaning to make it stable and readable.
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3. The vertical reading compass provides no relief from the requirement for magnetic correction compensation. Depending on the manufacturer and individual installation, the unit may even require supplem ental compensation to meet the maximum deviation requirements.
4. Vertical reading compasses are available on the aftermarket in both Technical Standard Order (TSO) and non-TSO versions. The altering technician must be aware of the limitations and certification responsibilities of the unit selected.
1105. TESTING, MARKING, AND PLACARDING.
a. Testing the Venturi Tube-Powered Systems. At normal in-flight cruise speed, or an accurately generated representative airflow, check the venturi tube-powered system to assure that the required vacuum is being supplied to the system.
b. Testing the Vacuum Pump-Powered System. When the system is powered by either engine or auxiliary driven vacuum pumps, check the system for proper output at their rated RPM. The output should be measured at the point of delivery to the instrument.
c. Testing of Altimeters and Static Systems.
Before performing an altimeter or static system test, determine that the system is free of contaminating materials such as dirt and water. With all instruments disconnected from the system, purge the plumbing with low pressure dry filtered air or nitrogen. Part 43, appendix E provides guidance and practical guidance on altimeter and static systems tests and inspections.
NOTE: Altimeter tests in accordance with part 43, appendix E must be accomplished by a properly certificated repair agency.
d. Testing Instrument Electrical Supply.
Subsequent to major repairs or alterations that results in addition, replacement, or rearrangement of instrument electrical circuitry, verify continuity and current availability before connecting the instruments. With th e individual instruments isolated from the power circuits, determine if the current available is in accordance with the manufacturer’s requirements.
e. Fuel, Oil, and Hydraulic Fluid Supply (Wet Instruments). Verify that the fluid transmission lines are free of residual material, dirt, and water before connecting them to the instruments. Purge them using a low pressure application of dry filtered air or nitrogen.
(1) On many CAR 3 and CAR 4a aircraft, common-sized fluid connections make it possible to incorrectly connect fluid lines to the wrong instruments. Be careful to assure delivery of fluids to the correct instrument.
NOTE: Fuel, Oil, and Hydraulic Fluid Supply (Wet Instruments), subparagraph (1) speaks to line fittings. Note that these wet instruments require a restricted fitting where the line attaches to the engine pressure outlet port. The restricted fitting is necessary to prevent fuel/oil/etc. from being sprayed into the cockpit during an instrument malfunction. There have been cases where this restricted fitting was removed and a normal fitting was installed, a fatal error for some personnel and airplane.
(2) Verify fluid pressures at both the source and instrument to confirm the absence of restrictions or blockage before connection to the instrument.
f. Fuel, Oil, and Hydraulic Fluid Supply (Dry Instruments). Use the same care exercised on the verification of security, source, and hygiene of wet instrument systems when checking dry systems.
(1) The primary difference is the delivery of the material to be measured to a transducer or Par 1104 Page 130
Par 1105 Page 131 pressure transmitter, rather than to the back of an instrument. Use of the wrong fluid can damage an instrument beyond repair.
(2) The additional difference between the two instrument systems is the attention required to provide the correct electrical circuitry and current between the transmitter and the instrument. (See subparagraph d.)
g. Instrument Markings and Placards.
(1) When additional instruments are installed, they must be appropriately marked. Refer to the applicable CAR and 14 CFR for specific instrument marking and placard requirements.
(2) Especially when the instrument panel was replaced, the altering individual should refer to the applicable ATC, TC, Group 2, Bulletin 7A, or TCDS of the aircraft to assure proper installation and placement of the required placards and limitations.
1106. ELECTRONIC DISPLAY INSTRU-
MENT SYSTEMS OR ELECTRONIC DISPLAY INDICATORS. The installation of these systems, regardless of the certification basis of the airplane into which they are being installed, are defined and specified under the provisions of § 23.1311.
a. Electrostatic Discharges.
(1) With new technologies come the requirements to learn new skills. While a certain amount of care and attention to detail while handling mechanical and electrical instruments has always been required, the hazards have been based on circumstances that could be seen and felt. With the glass cockpit, a different kind of hazard has surfaced as a real threat to airworthiness of instrumentation.
That threat is called Electrostatic Discharges (ESD).
(2) ESD is not a new phenomena. It has been with the industry a long time. However, with the advent of the electronic display instrument system and its multiple presentations, the problem can now disrupt entire display systems as a result of one event. Any unit containing electronic components such as diodes, transistors, integrated circuits, programmable read-only memory (ROM), ROM, and memory devices must be protected from ESD. The simple act of improperly carrying an electronic device across a room without adequate protection can render it unairworthy.
CAUTION: To prevent damage due to excessive electrostatic discharge, use proper gloves, finger cots, or grounding bracelets. Observe the standard procedures for handling equipment containing electrostatic sensitive devices or assemblies in accordance with the recommendations in the manufacturer’s maintenance instructions.
b. Electronic Display Instrument Systems Provide No Relief from the “Basic T”.
Section 23.1311(a)(5) requires systems using electronic displays to have independent backup systems to provide basic IMC reference in case electrical power is lost.
(1) Subparagraph 5 requires an independent magnetic indicator and either an independent secondary mechanical altimeter, airspeed indicator, and attitude instrument; or
(2) Individual electronic display indicators for altitude, airspeed, and attitude that are independent from the airplane’s primary electrical system.
(3) These secondary instruments may be installed in panel positions that are displaced from the primary positions as specified by § 23.1321(d), but must be located wher e they meet the pilot’s visibility requirements of § 23.1321(a).
c. Advisory Circular (AC) 23.1311-1, Installation of Electroni c Display in Part 23 Airplanes (as amended), paragraph 3.0, provides valuable interface references relating to the relationships between the CAR and 14 CFR
regulatory requirements installation of this technology to older and current technology in airplanes.
1107. ENVIRONMENTAL CONDITIONS.
a. General. The equipment environmental limits established by the manufacturer should be compatible with the operating environment of the airplane. When evaluation of the equipment installation, consider such factors as the maximum operating altitude of the airplane and whether the equipment is located within a temperature- and pressure-controlled area. Applicable methods for testing the performance characteristics of the equipment for specific environmental conditions are provided in the current edition of AC 21-16, RTCA, Inc. Document RTCA/DO-160E, Environmental Conditions and Test Procedures for Airborne Equipment. Either test or analysis, or both, ensures the compatibility between the operational environment and the environmental equipment category of the laboratory tests.
b. Temperature. An electronic system’s reliability is strongly related to the temperature of the solid-state components in the system. Component temperatures depend on the internal thermal design and external cooling. In evaluating the temperature requirements, consider the additional heat generated by the equipment, especially in an area where air flow is restricted. To determine if adequate cooling is provided, the evaluation must make maximum use of previous data from compatible installations. This will assist in limiting the ground and/or flight tests of those installations that cannot be verified by other means. When the equipment operating environment cannot be verified from previous experience or from evaluation of temperature values in that equipment location, a cooling test must be conducted.
c. Attitude Information. Attitude information should continue to be presented for a minimum of 30 minutes after the in-flight loss of cooling for the primary in strument when in the normal operating environment (temperature/ altitude). If proper perf ormance of the flight instrument functions is adversely affected due to loss of in-flight cooling, such conditions must be annunciated. Consider incorporation of an overtemperature shut-down of the system in case of cooling system failure. If such systems are used, AFM documentation should be established, requiring subsequent pilot actions. Additionally, applicable placards must be provided for pilot situational awareness of the critical condition. These actions should include procedures to allow possible recovery of a system that has had an overtemperature shutdown condition.
d. Annunciation. Annunciation of in-flight loss of cooling or fan monitors may not be required if shown by a safety analysis or test demonstration that a hazardous or catastrophic condition is not indicated. The safety analysis should consider the reliability of the fans, redundancies of functions, reversionary features (such as the ability to transfer critical functions), the enunciation of overtemperature and its response time, and the availability of other flight instrumentation. In some systems, cooling fans may be installed to improve the operating environment of the components and reduce the possibility of a failure condition or shutdown of the equipment. With supplementary installations, fan monitoring or additional temperature sensors may not be required. If cooling fans are needed to prevent a hazardous or catastrophic failure condition, installation of fan monitoring or other methods of determining the status of the cooling fan must be provided prior to flight.
1108. THRU 1199. RESERVED
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CHAPTER 12. CARGO TIEDOWN DEVICE INSTALLATIONS 1200. PURPOSE. This chapter provides data for making acceptable cargo tiedown device installations in non-pressurized areas of civil aircraft of 12,500 lbs gross weight or less. Engineering assistance is required for floor/attach fittings load analysis and material burn testing.
1201. HAZARDS AND WARNINGS.
a. Structural failure may occur if aircraft floor loading limits are exceeded.
b. Fire hazards may exist if materials do not meet flame resistant specifications.
c. Exceeding manufacturer’s weight and balance limitations create unsafe flight conditions.
1202. INSTALLATION CONSIDERATIONS.
a. Assure that the altered aircraft can be operated within the permissible weight and center of gravity (CG) ranges.
b. Determine that there will be unobstructed access to all equipment and controls essential to the proper operation of the aircraft, required emergency exits, and emergency equipment.
c. Use only materials that are at least flame resistant for covering of floors and webbing material. Refer to the a pplicable airworthiness standards for the aircraft involved to determine the required flame-resistant qualities. For aircraft in air taxi or other commercial operations, refer to the applicable operating rule for special requirements regarding fire protection, cargo bins, location of cargo with respect to passengers, cargo compartment, or aisle width.
1203. FABRICATION AND INSTALLATION.
a. Cargo Tiedown Devices.
(1) Cargo tiedown devices may be assembled from webbing, nets, rope, cables, fittings, or other material that conforms to a FAA-PMA, TSO, NAS, AN, or MIL-SPEC standards. Use snaps, hooks, clamps, buckles, or other acceptable fasteners rather than relying upon knots for securing cargo. Install tensioning devices or other acceptable means to provide a method of tightening and adjusting the restraint system to fit the cargoes to be carried.
(2) Provide covers or guards where necessary to prevent damage to or jamming of the aircraft's equipment, structure, or control cables.
(3) Straps and nets manufactured with MIL-SPEC webbing and thread must be evaluated to determine that the working load requirements are met. All tiedown assemblies are only as strong as the weakest component in the system, including the point of attachment.
NOTE: The owner/operator is responsible for ensuring that a basis of approval/acceptance is obtained for the tiedown device before use on aircraft. This is accomplished by providing substantiating engineering data of the devices being used, including aircraft floor load limits.
b. Structural Attachment. Commerciallyavailable seat tracks, rails, or other types of anchor plates may be used for structural attachment, provided they conform to a NAS, AN, or MIL- SPEC standard. This type of hardware permits a ready means of mounting a wide variety of quickdisconnect fittings for cargo tiedown. Typical examples of such fittings and their attachments are shown in Figures 12-1 through 12-5.
(1) When installing these fittings, reinforce the existing floorboards and/or other Par 1200 Page 133
adjacent structure to obtain the necessary load carrying capacity. Seat tracks installed longitudinally across lateral floor beams generally require fulllength support for adequate strength and rigidity between beam attach points (see Figure 12-4).
(2) Consider the inherent flexibility of the aircraft structure and install any reinforcement in a manner that will avoid localized stress concentrations in the structural members/areas. Give specific attention to the size, shape, and thickness of the reinforcement, fastener size and pattern, and the effects of any adhesives used.
(3) Fittings used for cargo tiedown attachment need not be substantiated by static tests if it can be shown that th e fitting's rated minimum breaking strength would not be exceeded by the applicable static test loads. Existing racks, rails, or other points used for attachment may be verified by static tests, analysis, or a written statement by the aircraft manufacturer attesting to its adequacy to withstand the necessary loads.
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