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
AC 43.13-1A 1972
DATUM - DATUM D *-D Cic
NOSE WHEEL TYPE AIRCRAFT
DATUM LOCATED FORWARD OF THE
MAIN WHEELS C.O. - D - (4X4)
TAIL WHEEL TYPE AIRCRAFT
DATUM LOCATED FORWARD OF THE
MAIN WHEELS C.G. = D +(RX L.)
-DATUM • DATUM *D~ G.S
NOSE WHEEL TYPE AIRCRAFT
DATUM LOCATED AFT OF THE MAIN WHEELS C.O.. - (D+ (D+ L)
TAIL WHEEL TYPE AIRCRAFT
DATUM LOCATED AFT OF THE MAIN WHEELS C.O.= -D (REL) CG • Distance from datum to center of gravity of the aircraft.
W - The weight of the aircraft at the time of weighing.
D = The horizontal distance measured from the datum to the main wheel weighing point, I • The horizontal distance measured from the main wheel weighing point to the nose or tall weighing point.
F = The weight at the nose weighing point.
R = The weight at the tall weighing point.
FIGURE 18.4.-Empty weight center-of-gravity formulas.
Par 657 Chap 13
Page 248 AC 43.13-1A 1972
DATUM (LEADING EDGE OF
WING ROOT SECTION)
COMPUTED C.G.
TARE 27 LBS.
LEFT SCALE 565 LBS. • SCALE READING WEIGHING POINT RIGHT SCALE 564 LES. 67LBS.
(@ MAIN WHEEL)
D, MEASURED - L, MEASURED -WEIGHING POINT (E TAIL WHEEL) TO FIND: EMPTY WEIGHT AND EMPTY WEIGHT CENTER OF GRAVITY Datum is the leading edge of the wing (from aircraft specification) (D) Actual measured horizontal distance from the main wheel weighing point f main wheel) to the Datum-- (L) Actual measured horizontal distance from the rear wheel weighing point (G rear wheel) to the main wheel weighing point-- --22211 SOLVING: EMPTY WEIGHT Weighing Scale Net Point Reading * Tare # Weight # Right 564 564 Left 565 565 Rear 67 27 40 Empty Weight (W) 1169 SOLVING: EMPTY WEIGHT CENTER OF GRAVITY Formula: C.G. = D + B x L = 9" + 40 × 222 = 3" + 7.6" = 10.6" 1169 Reference for formula, Figure 13,4 This case is shown properly entered on a sample weight and balance report form, Figure 13.17 FIGURE 18.5.-Empty weight and empty weight center of gravity-tail-wheel type aircraft.
Par 657 Chap 13
AC 43.13-1A Pago 249 1972
DATUM (LEADING EDGE OF
WING ROOT SECTION) COMPUTED •
NOSE SCALE 464 LBS. LEFT SCALE 620LB8., TARE 5LBS.
TARE IO LBS. RIGHT SOALE 609 L8S., TARE 5 LBS.
WEIGHING POINT - D, MEASURED (E NOSE WHEEL)
WEIGHING POINT L, MEASURED (2 MAIN WHEEL) TO FIND: EMPTY WEIGHT AND EMPTY WEIGHT CENTER OF GRAVITY Datum is the leading edge of the wing (from aircraft specification) (D) Actual measured horizontal distance from the main wheel weight point (Qmain wheel) to the Datum---- ---- 34.011 (L) Actual measured horizontal distance from the front wheel weighing point (C front wheel) to the main wheel weighing point--- -- 67.8"1 SOLVING: EMPTY WEIGHT Weighing Scale Tare # Net Point Reading * Weight Right 609 604 Leit 820 5 615 From 464 10 454 Empty Weight (W) 1678 SOLVING: EMPTY WEIGHT CENTER OF GRAVITY Formula: C.O. = D - 2&1 - 3411 - 456 2 87.8 - 3411 - 18.811 = 15.711 Reference for formula, Figure 13,4.
FioURE 18.8.-Empty welght and empty weight center-of-gravity-nose-wheel type aircraft.
Chap 13 Par 687
Page 250 AC 43.12-1A 1972 C.G. RANGE, THE LOADED C.G. MUST BE
WITHIN THE LIMITS SHOWN ON THE AIR-
CRAFT SPECIFICATION FOR ALL FLIGHT OPERATIONS, DATUM FORWARD LIMIT •- REARWARD LIMIT
- C.G.
FIGURE 18.7.-Operating center-of-gravity range.
ance report when used in lieu of the landing gear. Typical locations of the weighing points are shown in figure 18.8.
n. Zero fuel weight is the maximum permissible weight of a loaded aircraft (passengers, crew, cargo, etc.) less its fuel. All weights in excess of maximum zero fuel weight must consist of useable fuel.
o. Minimum Fuel. Minimum fuel for balance purposes is no more than the quantity of fuel necessary for one-half hour of operation at rated maximum continuous power, and is the maximum amount of fuel used in weight and balance computation when low fuel may adversely affect the most critical balance condition.
p. Full Oll. Full oil is the quantity of oil shown in the aircraft specifications as oil capacity.
Use full oil as the quantity of oil when making the loaded weight and balance computations.
q. Tare. Tare is the weight of chocks, blocks, stands, etc., used when weighing aircraft, and Par 657 is included in the scale readings. Tare is deducted from the scale reading at each respective weighing point where tare is involved to obtain the actual aircraft weight.
658. WEIGHING PROCEDURE. Accepted procedures when weighing an aircraft are:
a. Remove excessive dirt, grease, moisture, etc., from the aircraft before weighing.
b. Weigh the aircraft inside a closed building to prevent error in scale reading due to wind.
c. To determine the center of gravity, place the aircraft in a level flight attitude.
d. Have all items of equipment included in the certificated empty weight installed in the aircraft when weighing. These items of equipment are a part of the current weight and balance report. (See paragraph 662.)
•. Properly calibrate, zero, and use the scales in accordance with the scale manufacturer's instructions. The scales and suitable support for the aircraft, if necessary, are usually placed Chap 13