Page 332 AC 43.13-1A 1972 Section 4. OVERHAUL OF PROPELLERS 603, HUB AND HUB PARTS. Disassemble propellers submitted for overhaul and clean all hub parts in accordance with the manufacturer's recommendations. Make an inspection of the parts to determine that the critical dimensions are within the manufacturer's specified tolerances. Take particular care to check the 90° relation between shaft bore and blade socket centerline and track of the blade sockets, as these are the dimensions which are most likely to be affected by accidents. Reject and replace any hub which is sprung, worn, or damaged.
Stress risers such as cuts, nicks, or tool marks must be carefully stoned or the part rejected Carefully inspect splines and cone seats for signs of wear, and check splines with a single key "no-go" gauge made to plus 0.002 inch of the base drawing dimensions for spline land width. If the gauge enters more than 20 percent of the spline area, reject the hub. Cones and cone seats may show discoloration, pitting, and corrosion. Generally, corrosion and discolored spots may be removed by light lapping.
Pitting is not grounds for rejection, if 75 percent of the bearing area is not affected and the pitted areas are well dispersed about the cone bearing area. After cleaning, minutely inspect steel hubs and parts for cracks by the wet or dry magnetic particle method at every major overhaul period. It is not necessary to remove the plating or special external finish for this inspection unless so specified in the manufacturer's recommended practice.
Steel hubs which adapt fixed-pitch propellers to all taper crankshafts are susceptible to cracks along the keyway which often extend into the flange lightening holes. Carefully inspect these hubs by the magnetic particle method at engine overhaul periods. Any crack is cause for rejection.
604. PLATING FOR HUBS AND PARTS. Replate hubs and parts from which the plating has been removed after they have been satisfactorily inspected. All replating must be of the same material as the original plating and be done in accordance with the manufacturer's recommendations. Chrome or nickel plating is not an acceptable substitute for cadmium or zinc plating. The use of zinc chromate primer on the external surfaces followed by a coating of aluminum lacquer in lieu of cadmium plating is considered acceptable where recommended by the manufacturer. However, this type of finish will require replacement more frequently than cadmium plating.
605. INSPECTION OF ALUMINUM PROPELLERS AND BLADES. Carefully inspect aluminum propellers and blades for cracks and other injuries. A transverse (chordwise) crack or flaw of any size is cause for rejection. Refer any unusual condition or appearance revealed by these inspections to the manufacturer. Acceptable inspection methods are the acid-etching process, the anodizing process, the fluorescent penetrant process, or the dye penetrant process.
Where applicable, supplement one process with another. Etching is accomplished by immersing the blade in a warm 20 percent caustic soda solution and cleaning with a warm 20 percent nitric acid solution, with a warm water rinse between the caustic bath and the acid bath, and also a warm water rinse following the acid bath. Remove all effects of the etching by polishing. Maintain the caustic and acid solution at a temperature of from 160° F. to 180° F. Some blades incorporate parts made of steel and other materials, and the caustic soda and the nitric acid must not be allowed to come in contact with these parts. The blade surfaces are then examined with a magnifying glass of at least three-power. A crack will appear as a distinct black line. The fluorescent penetrant method is recommended as a supplement to the caustic etch for the inspection of Par 603 Chap 12
AC 43.13-1A 1972 the shanks (roots or hub ends) of adjustable pitch blades.
a. Suspected cracks or defects should be repeatedly local etched until their nature is determined. With a No. 00 sandpaper, or fine crocus cloth, clean and smooth the area containing the apparent crack. Apply a small quantity of caustic solution to the suspected area with a swab or brush. After the area is well darkened, thoroughly wipe it with a clean (dampened) cloth. Too much water may entirely remove the solution from a crack and spoil the test. If a crack, extending into the metal exists, it will appear as a dark line or mark, and by using a magnifying glass, small bubbles may be seen forming in the line or mark. Immediately upon completion of the final checks, remove all traces of the caustic solution by use of the nitric acid solution. Wash the blade thoroughly with clean (fresh) water.
b. The chromic acid anodizing process is superior to caustic etching for the detection of cracks and flaws and should therefore be used, whenever it is available, for general inspection of blades, for material defects, and for final checking of repairs performed during overhaul. Immerse the blades in the anodizing bath as far as possible, but all parts not made of Chap 12 Page 233 aluminum alloy must be kept out of the chromic acid bath or be separated from the blade by nonconductive wedges or hooks. Follow the anodizing treatment by a rinse in clear, cold, running water from 8 to 5 minutes, and dry the blades as quickly as possible, preferably with an air blast. Allow the dried blades to stand for at least 15 minutes before examination. Flaws (cold shuts or inclusions) will appear as fine black lines. Cracks will appear as brown stains caused by chromic acid bleeding out onto the surface. The blades may be sealed for improved corrosion resistance by immersing them in hot water (180° F. to 212° F.) for one-half hour. In no case, immerse the blades in hot water before the examination for cracks, since heat expands cracks and allows the chromic acid to be washed away.
606. ASSEMBLY. Accomplish the assembly of the propeller hub and blade in accordance with the manufacturer's recommendations. Replace clevis pins, bolts, and nuts which show wear or distortion. Never use cotter pins and safety wire a second time. The use of self-locking nuts is permissible only where originally used or approved by the manufacturer.
607-617. RESERVED.
Par 605
Page 224 AC 43.72-1A 1972 Section 5. ASSEMBLY OF PROPELLER TO ENGINE
618. FIXED-PITCH PROPELLERS. Loose hub bolts and bolts installed through the lightening holes in the integral hub flange of certain engine crankshafts cause the majority of the serious difficultles experienced with fixed-pitch propellers. Either of the conditions, if not corrected, will ultimately cause the loss of the propeller.
o. Loose hub bolts cause elongated boltholes and damage to the hub bolts. When not corrected, the bolts break off, or friction causes enough heat to affect the glue and chars the wood. After successive running, checks start at the boltholes. These checks are caused, or at least accentuated, by shrinkage of the wood due to the excessive heat generated. If allowed to progress, the propeller usually flies apart or catches on fire.
b. On some engines equipped with a crankshaft having an integral propeller hub flange, the outer edge of the lightening holes are at the same radius as the corresponding edge of the propeller hub boltholes. When inserting the bolts through the propeller, care must be exercised so that the bolts are inserted through the proper holes in the flange. Cases have been reported where the bolts were inserted through the larger lightening holes and, accordingly, the bolt nuts bore only on the outer edges of the lightening holes. In such cases, continuous running of the propeller may cause the boltheads or nuts to slip off the flange and through the large openings in the flange, resulting in the subsequent loss of the propeller.
c. Both the conditions discussed above are very easy to detect, and must be corrected immediately. In case the hub flange is integral with the crankshaft, frst ascertain the bolts are properly installed, then make the inspection for bolt tightness in the same manner as for any other propeller hub. Use an open-end wrench to determine hub bolt tightness; and, if the nuts can be turned, remove the cotter keys and retorque the nuts to the desired setting.
Tighten hub bolts, preferably with a torque wrench, to the recommended values which usually range from 15 to 24 foot-pounds. If no torque wrench is available, an ordinary socket wrench may be used. This socket wrench should have a 1-foot extension lever and the wrench pulled up with the recommended force 12 inches away from the center of the bolt which is being tightened. The tightening is best accomplished by tightening each bolt a litthe at a time, being sure to tighten alternate bolts which are diametrically opposite. Exercise care not to overtighten the hub bolts, thereby damaging the wood underneath the hub flanges. Avoid the practice of overtightening bolts to draw a propeller into track. Safety the nuts by means of cotter keys of proper size or heavy safety wire twisted between each nut.
A continuous length of single safety wire is not acceptable, as wire failure will result in all nuts becoming unsafetied.
619,-629. RESERVED.
Par 618 Chap 12
AC 43.13-18 Page 233 1972 Section 6. PRECAUTIONS
630. BLADE-TIP IDENTIFICATION. Many persons have been fatally injured walking into whirling propellers. Painting a warning strip on the propeller serves to reduce chances of such injuries. Cover approximately 4 inches of the propeller tips on both sides with an orange-yellow nonreflecting paint or lacquer. Open the drain holes in the metal tipping of wood blades after the tips have been painted.
631. GROUND-HANDLING PRACTICES, Wood propellers are especially susceptible to damage from improper handling. When moving an airplane, avoid bumping the propeller. The practice of pushing or pulling on a propeller blade to move an airplane should be avoided; it is extremely easy to impose forces on a blade in excess of those for which the blade is designed. It is continually necessary to ascertain that the glue joints are in good condition, and that the finish on the entire propeller will protect the propeller from absorbing moisture. Place twobladed wood propellers, whether on or off an airplane, in a horizontal position to prevent unbalance from moisture absorption. A good precaution is to cover the propeller with a well-fitted waterproof cover when not in use.
It is very important to protect the shank section of wood blades from moisture changes to prevent swelling and subsequent loosening in the metal sleeve. In the case of varnished blades, it is advisable to occasiona ly apply varnish around the shank at the junction or wood and metal. In the case of the plastic-covered blade, repair cement may be applied around the same joint.
632. COMPOSITION PROPELLER SERVICE LIMITS.
In certain cases where the blade has been manufactured from laminated planks of composition material, longitudinal cracks or splitting between laminations have been observed after several hundred hours of operation. Do not permit these cracks to progress beyond the limits established by the manufacturer.
633.-643. RESERVED.
Chap 12 Par 630
Page 236 AC 43.13-1A 1972 Section 7. MAINTENANCE OF ROTOR BLADES
644. GENERAL. The design, materials, construction, and performance requirements for rotorcraft main and tail rotor blades vary considerably from those used in the manufacture and operation of airplane propellers. Major dissimilarities also exist among the blades fabricated for the various makes and models of rotorcraft currently in operation. For these reasons, it is of utmost importance that the specific maintenance requirements of each rotorcraft manufacturer be consulted prior to performing any maintenance. In all instances, the restoration of blades will be determined by the repair limitations provided in the blade manufacturer's current service and maintenance manual data.
The information in this section may serve as a general guide for maintaining the airworthiness of rotorcraft blades.
a. Leading Edge Strip. Examine leading edge abrasion strips (A in figures 12.8, 12.9 and
12.10) for nicks, cuts, gouges, or worn spots resulting from abrasion. Damage of this type can usually be considered minor provided the injuries are not concentrated in one area, do a fed tho not penetrate the strip, or do not affect the aerodynamic performance of the blade.
rations or edge voids (B in Figure 12.8) are es- a pid by The limits for leading edge strip bond sepatablished by the blade manufacturer. When a separation is suspected, carefully examine the blade assembly and determine whether it is within the manufacturer's tolerances for continued operation. If it is permissible for the blade to remain in service, recheck that blade area at frequent intervals thereafter to assure that the separation is not progressing beyond allowable limits.
Inspect the rotor blade for evidence of paint deterioration in the area where the aft edge of the leading edge metal strip terminates. If the surface of the paint in this area is cracked, blistered, or missing, carefully examine the metal surface of the blade for evidence of electrolytic corrosion or pitting (C in Figure 12.8).
Treat minor cases of localized corrosion promptly by cleaning and coating the area with a preservative that will arrest further surface deterioration until the blade can be completely refinished. When the area of corrosion appears to be extensive or pitting is deep, remove the blade for complete evaluation of the damage.
b. Blade Surface. Examine the blade upper and lower surfaces between the leading edge strip and the trailing edge for evidence of cracks (B in figures 12.9 and 12.10). Cracks of any magnitude in this area are cause for blade removal if allowable crack limitations are not prescribed by the blade manufacturer.
Trailing edge cracks in blades having bonded trailing edges are also cause for removing the blade for service. Blades incorporating rivets along the trailing edge that have cracks at the edge (C in figures 12.9 and 12.10) which terminate at a rivet hole or cracks that extend between two adjacent rivets (D in figures 12.9 and 12.10) are not permissible and the blade should be removed from further service until repairs are made.
Inspect the adhesive bonding at the blade root end doublers (E in figures 12.9 and
12.10), skin to spar joint (A in figure 12.11), and the trailing edge (C in figure 12.11) for evidence of bond separation. This may initially appear as a hairline crack and any indication of a separation is cause for removing the blade from service to thoroughly examine and evaluate the condition prior to attempting repair.
Check the entire bonded (honeycomb core) area of the blade for evidence of bond separations or voids (B in figure 12.11). In many instances these can be detected by a difference in sound heard while the surface of the blade is being tapped with a light object such as a coin.
A dull or dead sound is indicative of a void or separation. The acceptability of voids or sepa- Par 644 Chap 12
AC 43.13-1A 1972 FIGURE 12.11.-Blade cross-section.
Par 644 Chap 12
AC 43.13-1A 1972 B FIGURE 12.12.-Blade retention and link attach points.
Par 644 Chap 12
AC 43.13-1A 1972 rations dependa upon their location and the size of the area involved. Consult the applicable blade manufacturer's service manual for the allowable limits.
Nore: The presence of water in the honeycomb core of a blade is cause for retiring a blade from service.
Examine the blade trim tabs for security of attachment and evidence of cracks in the area where they are bent (D in figure 12.8). When either of these conditions exists, take appropriate corrective action prior to approving the blade for return to service.
c. Main Spar or Beam. Inspect all of the exposed surfaces of blade spars for such damage as cuts, nicks, scratches, or erosion. When the blade leading edge is an integral part of an extruded spar or beam (D in figure 12.11), carefully examine the blade edge for wear resulting from abrasives. Use a penetrant method of inspection when checking the exposed surfaces of the metal spar or beam for cracks. If the spar is not exposed, an X-ray or comparable type of inspection is recommended. Any spar or beam crack is cause for removing a blade from service. Other types of spar damage must be evaluated, and if permissible, repaired in accordance with data provided in the applicable manufacturer's service manual.
d. Blade Retention. Examine the blade main retention components and drag (lead/lag) link attach points (A and B in figure 12.12) for evidence of looseness. If there are any indications of excessive play, remove the main retention and drag link attaching bolts or pins and check them for wear or stretch. Use a dye penetrant or magnetic particle method to inspect Chap 12 Page 241 the bolts or pins for cracks. Check the threads for evidence of necking or distortion. Examine the main and drag link bushings for wear or elongation. Carefully inspect the area around the bushing holes for evidence of cracks or delamination. Consult the manufacturer's overhaul or service manual to determine what main retention and drag link attachment wear tolerances are acceptable.
e. Blade Tip Assembly. Inspect the tip area of the blade for evidence of abrasive erosion. If this condition is readily visible and appears to be deep or covers an appreciable area, the blade should be removed from service in order to fully evaluate the degree of wear. Accomplish repairs in accordance with the manufacturer's recommendations if the wear limits for continued operation are exceeded.
Check the blade tip for loose rivets or bond separation. If either condition is evident, remove the blade from service and perform the necessary repairs. Any deformation of the blade tip must be corrected in the manner specifled by the manufacturer of the blade.
If a blade tip cap is installed, remove it and carefully examine the blade balance weights for security of attachment. Check the condition of the tip cap screws and anchor nuts prior to replacing the cap.
f. Blade Life. Most rotor blade manufacturers prescribe permissible limits for wear, repairs, and hours of operation for specific blade models. Close adherence to these limitations is recommended, to assure continued airworthiness and reliability of the blades.
645.-655. RESERVED.
Par 644
AC 43.13-1A Pago 243 1972 Chapter 13. WEIGHT AND BALANCE
656. GENERAL. The removal or addition of equipment results in changes to the center of gravity and empty weight of the aircraft, and the permissible useful load is affected accordingly. Investigate the effects of these changes, otherwise the aircraft flight characteristics may be adversely affected. Information on which to base the record of weight and balance changes to the alrcraft may be obtained from the pertinent aircraft specification, the prescribed aircraft operating limitations, airplane flight manual, and the aircraft weight and balance report.
Removal or addition of minor items of equipment such as nuts, bolts, rivets, washers, and similar standard parts of negligible weight on fixed-wing aircraft do not require a weight and balance check. Since rotorcraft are in general more critical with respect to control with changes in c.g. positions, the procedures and instructions in the particular model maintenance or flight manual should be followed.
657. TERMINOLOGY. The following terminology is used in the practical application of weight and balance control.
a. Maximum Weight. The maximum weight is the maximum authorized weight of the aircraft and its contents as listed in the specifications.
b. Empty Weight. The empty weight of an aircraft includes all operating equipment that has a fixed location and is actually installed in the aircraft. It includes the weight of the airframe, powerplant, required equipment, optional and special equipment, fixed ballast, full engine coolant, hydraulic fluid, and the fuel and oil as explained in paragraph 658 f and g.
Additional information regarding fluids which may be contained in the aircraft systems and which must be included in the empty weight will be indicated in the pertinent aircraft specifications whenever deemed necessary.
c. Useful Load. The useful load is the empty weight subtracted from the maximum weight of the aircraft. This load consists of the pilot, crew if applicable, maximum oil, fuel, passengers, and baggage, unless otherwise noted.
d. Weight Check. A weight check consists of checking the sum of the weights of all items of useful load against the authorized useful load (maximum weight less empty weight) of the aircrait.
e. Dotum. The datum is an imaginary vertical plane from which all horizontal measurements are taken for balance purposes with the aircraft in level flight attitude. The datum is indicated on most FAA Aircraft Specifications. On some of the older aircraft, where the datum is not indicated, any convenient datum may be selected. However, once the datum is selected, all moment arms and the location of the permissible c.g. range must be taken with reference to it. Examples of typical locations of the datum are shown in figure 18.1.
f. Arm (or Moment Arm), The arm is the horizontal distance in inches from the datum to the center of gravity of an item. The algebraic sign is plus (+), if measured aft of the datum, and minus (-) if measured forward of the datum.
Examples of plus and minus arms are shown in figure 18.2.
g. Moment. Moment is the product of a weight multiplied by its arm. The moment of an item about the datum is obtained by multiplying the weight of the item by its horizontal distance from the datum. A typical moment calculation ia given in figure 18.3.
h. Center of Gravity. The center of gravity is a point about which the nose-heavy and tailheavy moments are exactly equal in magnitude.
If the aircraft were suspended from there, it would have no tendency to pitch in either di- Chap 13 Par 656
Pago 244 AC 43.13-1A 1972 -DATUM
- DATUM
- DATUM
ANY CONVENIENT DISTANCE FIoURE 18.1.--Typical datum locationa.
Par 657 Chep 13
AC 43,13-1A Page 245 1972
F DATUM OIL COOLER 5LBS./ BAGGAGE 60LBS.
ARM 20" - ARM 60"-
MINUS ARM PLUS ARM FigURE 18.2.-Illustration of arm (or moment arm).
rection (nose up or down). The weight of the aircraft (or any object) may be assumed to be concentrated at its center of gravity. (See figure 13.8.)
I. Empty Weight Center of Gravity. The empty weight c.g. is the center of gravity of an aircraft in its empty weight condition, and is an essential part of the weight-and-balance record. Formulas for determining the center of gravity for tail and nosewheel type aircraft are given in figure 18.4. Typical examples of computing the empty weight and empty weight
c.g. for aircraft are shown in figures 13.5 and
18.6.
j. Empty Weight Center-of-Gravity Range, The empty weight center-of-gravity range is determined so that the empty weight c.g. limits will not be exceeded under standard specification loading arrangements. In cases where it is possible to load an aircraft in a manner not covered in the Aircraft Specification (i.e., extra tanks, extra seats, etc.), complete calculations as outlined in paragraph 661. The empty weight c.g. range, when applicable, is listed on the Aircraft Specifications. Calculation of Chap 13 empty weight c.g. is shown in figures 18.5 and
13.6.
k. Operating Center-of-Gravily Range. The operating c.g. range is the distance between the forward and rearward center of gravity limita indicated on the pertinent aircraft specification. These limits were determined as to the most forward and most rearward loaded c.g.
positions at which the aircraft meets the requirements of the Federal Aviation Regulations. The limits are indicated on the specification in either percent of mean aerodynamic chord or inches from the datum. The c.g. of the loaded aircraft must be within these limits at all times as illustrated in figure 18.7.
I. Mean Aerodynamic Chord (MAC), For weight and balance purposes it is used to locate the
c.g. range of the aircraft. The location and dimensions of the MAC will be found in the Aircraft Specifications, Aircraft Flight Manual, or the Aircraft Weight and Balance Report.
m. Weighing Point. If the c.g. location is determined by weighing, it is necessary to obtain horizontal measurements between the points on the scale at which the aircraft's weight is Par 657
Pao 246 AC 43.13-14 1972 DATUM
CENTER OF GRAVITY OR
POINT OF BALANCE The entire aircraft weight may be considered to be concentrated at the center of gravity. Therefore, the moment of the aircraft about the datum is the weight of the aircraft times the horizontal distance between the C.G. and the datum.
Example: If the weight of this airplane is 2000 Ibs. and the arm from the datum to the center of gravity is 16 inches, the moment of the aircraft about the datum is 2000 × 16 or 32,000 in. Ibs.
FIGURE 18.8.-Example of moment computation.
concentrated. If weighed, using scales under the landing gear tires, a vertical line passing through the centerline of the axle will locate the point on the scale at which the weight is concentrated. This point is called the "weigh- Far 657 ing point." Other structural locations capable of supporting the aircraft, such as jack pads on the main spar, may also be used if the aircraft weight is resting on the jack pads. Indicate these points clearly in the weight and bal- Chap 13