FIGURE 10-17. SAMPLE WEIGHT AND BALANCE REPORT TO DETERMINE
EMPTY WEIGHT CENTER OF GRAVITY MAKE MA-700 MODEL A SERIAL # 0000 REGISTRATION # N1234 DATOM IS_ leading edge of wing
COMPUTE AS FOLLOWS IF AIRCRAFT WEIGHED
1. Leveling means: level top longeron between front and rear seats.
2. Main wheel weighing point is located - "FORWARD) (+ 3" AFT) of datum.
3. Actual measured distance from the main weight point centerline to the tail (or nose) point centerline 222 "
4. Oil over and above "ZERO" tank reading = (a. - Gals.) (b.- Lbs.) (c. - In.)
ACTUAL EMPTY WEIGHT Weight Point Scale Reading Tare = Net Weight
5. Right 564 564
6. Left 565 565
7. Tail 67 27 40
8. Nose
9. Total Net Weight 1169
CENTER OF GRAVITY AS WEIGHED
10. C.G. relative to main wheel weighing point:
(a) Tail wheel airc. (Item 3, 222), x (Item 7, 40 ) = C.G.
(Item 9, 1169)
(b) Nose wheel airc. (Item 3 ---) x (Item 8 ---- ) = C.G.
(Item 9 --- )
11. C.G. relative to datum:
(a) Tail wheel airc. (Item 10a, + 7.6) added to (Item 2, + 3) +10.6" = C.G.
(b) Nose wheel airc. (Item 10b. added to (Item 2, ) = = C.G.
COMPUTE IF AIRCRAFT WEIGHED WITH OIL (Item 4) Weight Arm Moment Aircraft (9) (11) Less Oil (4b) (4c) Empty Totals (a) (b)
(b) -- -= (c) = Empty weight C.G.
12. (a)
REPAIR AGENCY DATE Name Number Par 10-23 Page 10-22
FIGURE 10-18. SAMPLE WEIGHT AND BALANCE REPORT INCLUDING AN
EQUIPMENT CHANGE FOR AIRCRAFT FULLY LOADED
EQUIPMENT LIST *Required or Optional Item Numbers as Shown in Aircraft Specification 105 106 201 202 302(a) 303 401(a) 402 Special Equipment Item Make Model Weight Arm 3 Flares 1-1/2 Min XYZ 03 25# 105" Enter above those items included in the empty weight.
WEIGHT AND BALANCE EXTREME CONDITIONS Approved fid limit 8.5" Approved max. weight 2100# Approved aft limit 21.9" Item FORWARD CHECK REARWARD CHECK Weight Arm = Airo. Empty + 12391 + 1169 (9 or 12a) (11 or 12c) (9 or 12a) (11 or 12c) Oil + 17 - 49 - 833 Pilot + 170 + 16 + 2720 Fuel + 83 + 240 + 22 + 5280 Passenger (s + 340 + 48 + 16320 Baggage + 100 + 75.5 + 7550 TOTAL + 1439 = TW + 16104 = TM + 2036 = TW +43428=IM IM = 16104 = +11.2" = IM = 43428 = +21.3" = TW 1439 TW 2036 Most Forward C.G. location Most Rearward C.G. location
LOADING SCHEDULE Gallons Number of Pounds of of Fuel Passengers Baggage 40 2(R) 100 The above includes pilot and capacity oil
EQUIPMENT CHANGE Computing New C.G.
Item, Make, and Model* Weight Arm Moment Airc. Empty + 1169 + 10.6 + 12391 (9 or 12a) (11 or 12c) 204 added + 6 - 1 - 6 302(b) added + 29 + 13 + 377 302(a) removed - 24 + 29 + 696 303 removed - 1 + 4 - 4 NET TOTALS - 1179 = NW + 13454 = NM NM = 13454= +11.4" = New C.G.
NW 1179 *ITEM NUMBERS WHEN LISTED IN PERTINENT AIRCRAFT SPECIFICATION MAY BE USED IN LIEU OF "ITEM, MAKE, AND MODEL".
PREPARED BY DATE_ Par 10-23 Page 10-23
10-24. INSTALLATION OF BALLAST. Ballast is sometimes permanently installed for CG balance purposes as a result of installation or removal of equipment items and is not used to correct a nose-up or nose-down tendency of an aircraft. It is usually located as far aft or as far forward as possible in order to bring the CG position within acceptable limits with a minimum of weight increase. Permanent ballast is often lead plate wrapped around and bolted to the fuselage primary structure (e.g., tail-post, longerons, or bulkhead members). Permanent ballast invariably constitutes a concentrated load; therefore, the strength of the local structure and the attachment of the ballast thereto should be investigated for the design loading conditions pertinent to that particular aircraft. Placard permanent ballast with "Permanent ballast - do not remove." It is not desirable to install permanent ballast by pouring melted lead into the tail-post or longerons due to difficulties that may be encountered in subsequent welding repair operations. It should be noted that the installation of permanent ballast results in an increase of aircraft empty weight. See Figure 10-19 for ballast computation. The local strength of the compartment in which the ballast is carried and the effect of the ballast on aircraft W&B should be investigated when disposable ballast is carried.
FIGURE 10-19. PERMANENT BALLAST COMPUTATION FORMULA Desired C.G. location Actual C.G. location Location of ballast to D - Distance in inches desired to move C.G. of airplane. be added W - Weight of airplane as loaded.
X - Distance in inches from point where ballast is to be installed, to the desired location of the new C.G.
B - Weight of ballast required in pounds.
B = D × W Compute the new C.G. of the aircraft with ballast installed.
NOTE: If greater accuracy is desired, repeat the entire formula using the NEW aircraft weight and the NEW C.G. in the second operation.
Par 10-24 Page 10-24
Par 10-25 Page 10-25
10-25. LOADING SCHEDULE. The loading schedule should be kept with the aircraft and form a part of the AFM. It includes instructions on the proper load distribution such as filling of fuel and oil tanks, passenger seating, restrictions of passenger movement, and distribution of cargo.
a. Other means of determining safe loading conditions such as the use of a graphical index and load adjuster are acceptable and may be used in lieu of the information in paragraph 10-20.
b. Compute a separate loading condition when the aircraft is to be loaded in other than the specified conditions shown in the loading schedule.
10-26.–10-34. [RESERVED.]
Par 11-1 Page 11-1 CHAPTER 11. AIRCRAFT ELECTRICAL SYSTEMS SECTION 1. INSPECTION AND CARE OF ELECTRICAL SYSTEMS
11-1. GENERAL. The term “electrical system” as used in this AC means those parts of the aircraft that generate, distribute, and use electrical energy, including their support and attachments. The satisfactory performance of an aircraft is dependent upon the continued reliability of the electrical system. Damaged wiring or equipment in an aircraft, regardless of how minor it may appear to be, cannot be tolerated. Reliability of the system is proportional to the amount of maintenance received and the knowledge of those who perform such maintenance.
It is, therefore, important that maintenance be accomplished using the best techniques and practices to minimize the possibility of failure. This chapter is not intended to supersede or replace any government specification or specific manufacturer’s instruction regarding electrical system inspection and repair.
11-2. INSPECTION AND OPERATION CHECKS. Inspect equipment, electrical assemblies, and wiring installations for damage, general condition, and proper functioning to ensure the continued satisfactory operation of the electrical system. Adjust, repair, overhaul, and test electrical equipment and systems in accordance with the recommendations and procedures in the aircraft and/or component manufacturer’s maintenance instructions. Replace components of the electrical system that are damaged or defective with identical parts, with aircraft manufacturer’s approved equipment, or its equivalent to the original in operating characteristics, mechanical strength, and environmental specifications. A list of suggested problems to look for and checks (see Appendix 1, Glossary, for a description of the check types) to be performed are:
a. Damaged, discolored, or overheated equipment, connections, wiring, and installations.
b. Excessive heat or discoloration at high current carrying connections.
c. Misalignment of electrically driven equipment.
d. Poor electrical bonding (broken, disconnected, or corroded bonding strap) and grounding, including evidence of corrosion.
e. Dirty equipment and connections.
f. Improper, broken, inadequately supported wiring and conduit, loose connections of terminals, and loose ferrules.
g. Poor mechanical or cold solder joints.
h. Condition of circuit breaker and fuses.
i. Insufficient clearance between exposed current carrying parts and ground or poor insulation of exposed terminals.
j. Broken or missing safety wire, broken bundle lacing, cotter pins, etc.
Par 11-2 Page 11-2
k. Operational check of electrically operated equipment such as motors, inverters, generators, batteries, lights, protective devices, etc.
l. Ensure that ventilation and cooling air passages are clear and unobstructed.
m. Voltage check of electrical system with portable precision voltmeter.
n. Condition of electric lamps.
o. Missing safety shields on exposed high-voltage terminals (i.e., 115/200V alternating current (AC)).
11-3. FUNCTIONAL CHECK OF STAND-BY OR EMERGENCY EQUIPMENT. An aircraft should have functional tests performed at regular intervals as prescribed by the manufacturer. The inspections or functional check periods should be clearly stated in the aircraft maintenance manual, along with the overhaul intervals.
11-4. CLEANING AND PRESERVATION. Annual cleaning of electrical equipment to remove dust, dirt, and grime is recommended. Suitable solvents or fine abrasives that will not score the surface or remove the plating may be used to clean the terminals and mating surfaces if they are corroded or dirty. Only cleaning agents that do not leave any type of residue must be used. Components must be cleaned and preserved in accordance with the aircraft handbooks or manufacturer’s instructions. Avoid using emery cloth to polish commutators or slip rings because particles may cause shorting and burning. Be sure that protective finishes are not scored or damaged when cleaning. Ensure that metal-to-metal electrically bonded surfaces are treated at the interface with a suitable anti-corrosive conductive coating, and that the joint is sealed around the edges by restoring the original primer and paint finish. Connections that must withstand a highly corrosive environment may be encapsulated with an approved sealant in order to prevent corrosion.
CAUTION: Tu rn power off before cleaning.
11-5. BATTERY ELECTROLYTE CORROSION. Corrosion found on or near lead-acid batteries can be removed mechanically with a stiff bristle brush and then chemically neutralized with a 10-percent sodium bicarbonate and water solution. For Nickel Cadmium (NiCad) batteries, a 3-percent solution of acetic acid can be used to neutralize the electrolyte. After neutralizing, the battery should be washed with clean water and thoroughly dried.
11-6. ADJUSTMENT AND REPAIR. Accomplish adjustments to items of equipment such as regulators, alternators, generators, contactors, control devices, inverters, and relays at a location outside the aircraft, and on a test stand or test bench where all necessary instruments and test equipment are at hand. Follow the adjustment and repair procedures outlined by the equipment or aircraft manufacturer. Replacement or repair must be accomplished as a part of routine maintenance. Adjustment of a replacement voltage regulator is likely since there will always be a difference in impedance between the manufacturer’s test equipment and the aircraft’s electrical system.
Par 11-7 Page 11-3
11-7. INSULATION OF ELECTRICAL EQUIPMENT. In some cases, electrical equipment is connected into a heavy current circuit, perhaps as a control device or relay. Such equipment is normally insulated from the mounting structure since grounding the frame of the equipment may result in a serious ground fault in the event of equipment internal failure. Stranded 18 or 20 American Wire Gauge (AWG) wire should be used as a grounding strap to avoid shock hazard to equipment and personnel. If the end connection is used for shock hazard, the ground wire must be large enough to carry the highest possible current (0.1 to 0.2 ohms max.).
11-8. BUS BARS. Annually check bus bars for general condition, cleanliness, and security of all attachments and terminals. Grease, corrosion, or dirt on any electrical junction may cause the connections to overheat and eventually fail. Bus bars that exhibit corrosion, even in limited amounts, should be disassembled, cleaned and brightened, and reinstalled.
11-
9. ASTM INTERNATIONAL. The Federal Aviation Administration (FAA) worked in conjunction with ASTM International to initiate voluntary consensus standards. The FAA reviewed the consensus standards for compliance with the regulatory requirements and published FAA-acceptance of the standards in the Federal Register (FR) as an acceptable means of compliance to 14 CFR part 23 sections concerning electrical wiring systems. Any normal, utility, acrobatic, and commuter aircraft issued an airworthiness certificate, which has been designed, manufactured, and maintained in accordance with accepted ASTM consensus standards, provides the public with the appropriate level of safety established under the regulations. These standards are not intended to replace any instructions for continued airworthiness (ICA) published by the aircraft manufacturer or type design holder. The consensus standards may be used unless the FAA publishes a specific notification otherwise. The FAA maintains a listing of accepted consensus standards on the FAA website (www.faa.gov). Examples of consensus standards accepted by the FAA are shown in Table 11-1.
TAB LE 11-1. ASTM CONSENSUS STANDARDS Standard Designation Title Published By Federal Register Notice Organization Committee Date FR Page F2490-05 (2013) Standard Guide for Aircraft Electrical Load and Power Source Capacity Analysis ASTM International F39 13-Mar-2006 71 FR 12771 F2639-07€1 Standard Practice for Design, Alteration, and Certification of Airplane Electrical Wiring Systems ASTM International F39 23-Sep-2010 75 FR 58016 F2696-14 Standard Practice for Inspection of Aircraft Electrical Wiring Systems ASTM International F39 30-Dec-2014 79 FR 78553 F2799-14 Standard Practice for Maintenance of Aircraft Electrical Wiring Systems ASTM International F39 30-Dec-2014 79 FR 78553
Par 11-9 Page 11-4 NOTE: A standard is issued under a fixed designation (i.e., F2696-08); the number immediately following the designation indicates the year of original adoption or, in the case of revision, the year of last revision. A number in parentheses indicates the year of last reapproval. A reapproval indicates a 2-year review cycle completed with no technical changes. A superscript epsilon (€) indicates an editorial change since the last revision or reapproval.
A notice of availability will only be issued for new or revised standards.
Reapproved standards issued with no technical changes or standards issued with editorial changes only (i.e., superscript epsilon (€)) are considered accepted by the FAA without need for a notice of availability.
11-10.–11-14. [RESERVED.]
Par 11-15 Page 11-5 SECTION 2. STORAGE BATTERIES
11-15. GENERAL.
a. Aircraft batteries may be used for many functions, e.g., ground power, emergency power, improving DC bus stability, and fault-clearing. Most small private aircraft use lead-acid batteries. Most commercial and military aircraft use NiCad batteries. However, other types are becoming available such as gel cell and sealed lead-acid batteries. The battery best suited for a particular application will depend on the relative importance of several characteristics, such as weight, cost, volume, service or shelf life, discharge rate, maintenance, and charging rate. Any change of battery type may be considered a major alteration.
b. Storage batteries are usually identified by the material used for the plates. All battery types possess different characteristics and, therefore, must be maintained in accordance with the manufacturer’s recommendations.
WARNING: It is extremely dangerous to store or service lead-acid and NiCad batteries in the same area. Introduction of acid electrolytes into alkaline electrolyte will destroy the NiCad and vice-versa.
11-16. BATTERY CHARGING. Operation of storage batteries beyond their ambient temperature or charging voltage limits can result in excessive cell temperatures leading to electrolyte boiling, rapid deterioration of the cells, and battery failure. The relationship between maximum charging voltage and the number of cells in the battery is also significant. This will determine (for a given ambient temperature and state of charge) the rate at which energy is absorbed as heat within the battery. For lead-acid batteries, the voltage per cell must not exceed
2.35 volts. In the case of NiCad batteries, the charging voltage limit varies with design and construction. Values of 1.4 and 1.5 volts per cell are generally used. In all cases, follow the recommendations of the battery manufacturer.
11-17. BAT TERY FREEZING. Discharged lead-acid batteries exposed to cold temperatures are subject to plate damage due to freezing of the electrolyte. To prevent freezing damage, maintain each cell’s specific gravity at 1.275, or for sealed lead-acid batteries check “open” circuit voltage. (See Table 11-2.) NiCad battery electrolyte is not as susceptible to freezing because no appreciable chemical change takes place between the charged and discharged states.
However, the electrolyte will freeze at approximately minus 75 °F.
NOTE: Only a load check will determine overall battery condition.
Par 11-17 Page 11-6 TABLE 11-2. LEAD-ACID BATTERY ELECTROLYTE FREEZING POINTS Specific Freeze Point State of Charge (SOC) for Sealed Lead-Acid Batteries at 70 ° Gravity C. F. SOC 12 volt 24 volt
1.300 -70 -95 100% 12.9 25.8
1.275 -62 -80 75% 12.7 25.4
1.250 -52 -62 50% 12.4 24.8
1.225 -37 -35 25% 12.0 24.0
1.200 -26 -16
1.175 -20 -4
1.150 -15 +5
1.125 -10 +13
1.100 -8 +19
11-18. TEMPERATURE CORRECTION. U.S.-manufactured lead-acid batteries are considered fully charged when the specific gravity reading is between 1.275 and 1.300. A 1/3-discharged battery reads about 1.240 and a 2/3 discharged battery will show a specific gravity reading of about 1.200, when tested by a hydrometer and the electrolyte temperature is 80 °F. However, to determine precise specific gravity readings, a temperature correction (see Table 11-3) should be applied to the hydrometer indication. As an example, with a hydrometer reading of 1.260 and the temperature of the electrolyte at 40 °F, the corrected specific gravity reading of the electrolyte is 1.244.
Par 11-18 Page 11-7 TABLE 11-3. SULFURIC ACID TEMPERATURE CORRECTION Electrolyte Temperature Points to be Subtracted or Added to Specific Gravity Readings °C °F 60 140 +24 55 130 +20 49 120 +16 43 110 +12 38 100 +8 33 90 +4 27 80 0 23 70 -4 15 60 -8 10 50 -12 5 40 -16 -2 30 -20 -7 20 -24 -13 10 -28 -18 0 -32 -23 -10 -36 -28 -20 -40 -35 -30 -44
11-19. BATTERY MAINTENANCE. Battery inspection and maintenance procedures vary with the type of chemical technology and the type of physical construction. Always follow the battery manufacturer’s approved procedures. Battery performance at any time in a given application will depend upon the battery’s age, state of health, state of charge, and mechanical integrity.
a. Age. To determine the life and age of the battery, record the install date of the battery on the battery. During normal battery maintenance, battery age must be documented either in the aircraft maintenance log or in the shop maintenance log.
b. State of Health. Lead-acid battery state of health may be determined by duration of service interval (in the case of vented batteries), by environmental factors (such as excessive heat or cold), and by observed electrolyte leakage (as evidenced by corrosion of wiring and connectors or accumulation of powdered salts). If the battery needs to be refilled often, with no evidence of external leakage, this may indicate a poor state of the battery, the battery charging system, or an over charge condition.
Par 11-18 Page 11-8
(1) Use a hydrometer to determine the specific gravity of the battery electrolyte, which is the weight of the electrolyte compared to the weight of pure water.
(2) Take care to ensure the electrolyte is returned to the cell from which it was extracted. When a specific gravity difference of 0.050 or more exists between cells of a battery, the battery is approaching the end of its useful life and replacement should be considered.
Electrolyte level may be adjusted by the addition of distilled water.
c. State of Charge. Battery state of charge will be determined by the cumulative effect of charging and discharging the battery. In a normal electrical charging system, the battery’s generator or alternator restores a battery to full charge during a flight of 1 hour to 90 minutes.
d. Mechanical Integrity. Proper mechanical integrity involves the absence of any physical damage as well as assurance that hardware is correctly installed and the battery is properly connected. Battery and battery compartment venting system tubes, nipples, and attachments, when required, provide a means of avoiding the potential buildup of explosive gases, and should be checked periodically to ensure that they are securely connected and oriented in accordance with the maintenance manual’s installation procedures. Always follow procedures approved for the specific aircraft and battery system to ensure that the battery system is capable of delivering specified performance.
e. B attery and Charger Characteristics. The following information is provided to acquaint the user with characteristics of the more common aircraft battery and battery charger types. Products may vary from these descriptions due to different applications of available technology. Consult the manufacturer for specific performance data.
NOTE: Under no circumstances connect a lead-acid battery to a charger, unless properly serviced.
(1) Lead-acid vented batteries have a 2-volt nominal cell voltage. Batteries are constructed so that individual cells cannot be removed. Occasional addition of water is required to replace water loss due to overcharging in normal service. Batteries that become fully discharged may not accept recharge.
(2) Lead-acid sealed batteries are similar in most respects to lead-acid vented batteries, but do not require the addition of water.
(3) The lead-acid battery is economical and has extensive application, but is heavier than an equivalent performance battery of another type. The battery is capable of a high rate of discharge and low-temperature performance. However, maintaining a high rate of discharge for a period of time usually warps the cell plates, shorting out the battery. Its electrolyte has a moderate specific gravity, and state of charge can be checked with a hydrometer.
(4) Do not use high- amperage automotive battery chargers to charge aircraft batteries.
(5) NiCad vented batteries have a 1.2-volt nominal cell voltage. Occasional addition of distilled water is required to replace water loss due to overcharging in normal service. Cause of failure is usually shorting or weakening of a cell. After replacing the bad cell with a good cell,