Par 1-29 Page 1-13 in the finish, fastener holes, inspection/access openings, control system openings, drain holes, and the interfaces of metal fittings and the wood structure). The mechanic should also look for evidence of swelling or warpage of the aircraft’s wood structure, which would indicate underlying damage or decay. Particular attention should be paid to the wood structure immediately beneath the upper surfaces, especially under areas that are finished in dark colors, for signs of deteriorating adhesives. Cracks in wood spars are often hidden under metal fittings or metal rib flanges and leading edge skins. Any time a reinforcement plate exists that is not feathered out on its ends, a stress riser exists at the ends of the plate. A failure of the primary structure can be expected to occur at this point.
b. Tapping the wood structure with a light plastic hammer or screwdriver handle should produce a sharp solid report. If the suspect area sounds hollow and soft, further inspection is warranted by the following methods.
c. Probe the area in question, if accessible, with a sharp metal tool. The wood structure should be solid and firm. If the suspect area feels soft and mushy the mechanic should assume that the area is rotted. Disassembly of the structure is warranted at this point.
d. Prying the area of a bond joint will reveal any mechanical separation of the joint. If the mechanic detects any relative movement between two adjacent wood members, a failure of the bond is evident. Any loose fittings should arouse the mechanic’s suspicion, and the fittings should be removed to check for elongated bolt holes. Disassembly is warranted for further inspection.
e. Odor is an important indicator of possible deterioration. During the initial inspection, as the access panels are being removed from the structure, the mechanic should be aware of any areas that smell musty or moldy. These odors are indicative of the presence of moisture and associated fungal growth and decay.
f. Visual inspection requires looking at the wood structure both externally and internally for visual signs of decay or physical damage. Any accumulations of dirt, bird nests, or rodent nests are likely places to hold moisture and promote decay.
(1) The mechanic should remove any such accumulations that are found and inspect the area for signs of decay. Decay will appear as a dark discoloration or gray stains running along the grain and often a swelling of the wood member if still wet. Fittings will be imbedded in the wood instead of flush.
(2) Highly suspected structurally damaged areas are shown in Figure 1-3. A list of most likely areas to incur structural damage include the following:
(a) Check front and rear spars for compression cracks adjacent to the plywood reinforcing plates, where the lift struts attach, and at the rib attach points on either side of the strut attach points. Triple-check these areas and the spar to fuselage attach points for cracks if the wingtip has contacted the ground, a hangar wall, etc.
(b) Check all metal fittings which attach to wooden structure for looseness, corrosion, cracks, or warps. Areas of particular interest are strut attach fittings, spar butt fittings,
Par 1-29 Page 1-14 aileron and flap hinges, jury strut fittings, compression struts, pulley brackets, and any landing gear fittings.
(c) Check front and rear spars for longitudinal cracks at the ends of the plywood reinforcement plates where the lift struts attach. Triple-check this area if the wing has encountered any kind of ground strike.
(d) Check ribs on either side of strut attach points for missing or loose rib-to-spar attach nails.
(e) Check ribs on either side of strut attach points for cracks where the cap strips pass over and under the spars.
(f) Check for cracked leading edge skin and/or failed nose ribs in the area directly in front of the jury strut.
(g) Check the brackets which attach the struts to the spars for cracks.
(h) Check the aileron, flap hinge, and hinge brackets for cracks and loose or missing rivets.
(i) Check all exposed end grain wood, particularly the spar butts, for cracking or checking. Checking, or splitting, of wood spar butts is common on aircraft based in arid areas.
(j) Also check for any cracks that indicate a bond line failure or structural failure of the wood member. Any evidence of movement of fittings, bushings, or fasteners should be cause for concern, and further inspection is warranted. Splits in fabric covering the plywood, especially on upper surfaces exposed to ultraviolet light and water, dictate that the mechanic remove the fabric around the split so the underlying plywood may be inspected for physical damage or decay. When removing metal fasteners from wood, check for evidence of corrosion. Any corrosion present indicates the presence of moisture and the strong probability of decay in the adjoining wood structure.
(k) Any wooden member that has been overstressed is subject to compression failure (e.g., ground loop). Compression cracking and failure of the wood spars in certain aircraft are a continuing problem. Compression failures are defined as failure of wood fibers on a plane perpendicular to the wood fiber’s longitudinal axis. If undetected, compression failures may result in structural failure of the wing during flight. Compression cracks have been found emanating from the upper surfaces of the wing spars and progressing downward.
(l) The usual locations for cracks have been the front spar at both ends of the reinforcement plate for the lift strut and the front spar rib attach points, both inboard and outboard of the spar reinforcement plate; and the rear spar lift strut and rib attach points. An inspection of both the front and rear spars for compression cracks is recommended.
(m)The two areas where it is possible to identify a compression crack are on the face and top surface of the spar. Using a borescope through existing inspection holes is one method of inspection. An alternate method is to cut inspection holes in the skin. If inspection
Par 1-29 Page 1-15 holes are cut, they should be made on the aft side of the front spar and the forward side of the rear spar. This will allow the fabric to be peeled away from the spar. Longitudinal cracks may also be detected during this inspection. Loose or missing rib nails may indicate further damage and should be thoroughly investigated. The mechanic may shine a light, at a low angle and parallel with the grain, in the area of the member subjected to the compression load. An area of grain waviness would indicate a potential compression failure. In all cases the manufacturer’s inspection data should be followed.
FIGURE 1-3. LIKELY AREAS TO INCUR STRUCTURAL DAMAGE
g.
Moisture meters are effective tools for detection of excessive moisture content in wood members. An instrument such as this allows the mechanic to insert a probe into the wood member and read its moisture content directly off the meter. A correction chart usually accompanies the instrument to correct for temperature and species of wood. Any reading over 20 percent indicates the probability of fungus growth in the member. Moisture content of the wood should be 8–16 percent, preferably in the 10–12 percent range (this range is during inspection). Where plywood skin covers the spar and the spar would be inaccessible without removing the skin, the moisture meter probe can be inserted through the plywood skin and into the spar to check the moisture content of the spar. The small holes made by the probe are easily sealed.
h. Destructive testing of sample bonded joints whenever a new bond joint is made, a sample joint should be made with the adhesive from the same batch used on the repair and scraps of wood left over from the repair. After curing, the sample joint should be destructively tested to ensure proper bonding of the two wood pieces. Any failure in the bond line indicates a cohesive failure of the adhesive. Any failure along the bond line indicates an adhesive failure, which is indicative of poor bonding. The ideal situation is when wood fibers are observed on both sides of the fracture surface. This indicates a failure in the wood, and indicates the bond joint is actually stronger than the wood.
1-30.–1-35. [RESERVED.]
Par 1-36 Page 1-16 SECTION 4. REPAIRS
1-36. GENERAL. The basic standard for any aircraft repair is that the repaired structure must be as strong as the original structure and be equivalent to the original in rigidity and aerodynamic shape. Repairs should be made in accordance with manufacturer specifications whenever such data is available.
1-37. REPLACEMENT OF DRAIN HOLES AND SKIN STIFFENERS. Whenever repairs are made that require replacing a portion that includes drain holes, skin stiffeners, or any other items, the repaired portion must be provided with similar drain holes, skin stiffeners, or items of the same dimensions in the same location. Additional drain holes may be required if reinforcement under a skin repair interferes with water flow to existing drain holes. Make any additional drain holes the same diameter as originals, usually 1/4 inch.
1-38. CONTROL SURFACE FLUTTER PRECAUTIONS. When repairing or refinishing control surfaces, especially on high performance airplanes, care must be exercised that the repairs do not involve the addition of weight aft of the hinge line. Such a procedure may adversely affect the balance of the surface to a degree that could induce flutter. As a general rule, it will be necessary to repair control surfaces in such a manner that the structure is identical to the original, and that the stiffness, weight distribution, and mass balance are not affected in any way. Consult the aircraft maintenance manual or seek manufacturer’s direction for specific requirements on checking control surface balance after repair and refinishing of any control surface.
1-39. SCARF JOINTS. The scarf joint is the most satisfactory method of making an end joint between two solid wood members. Cut both parts accurately. The strength of the joints depends upon good joint design and a thin, uniform bond line. Make the scarf cut in the general direction of the grain slope as shown in Figure 1-4.
Par 1-39 Page 1-17 FIGURE 1-4. CONSIDERATION OF GRAIN DIRECTION WHEN MAKING
SCARF JOINTS 1 -40. SPLICING OF SPARS. Unless otherwise specified by the manufacturer, a spar may be spliced at any point except under the wing attachment fittings, landing gear fittings, engine mount fittings, or lift and interplane strut fittings. These fittings may not overlap any part of the splice. A spar splice repair should not be made adjacent to a previous splice or adjacent to a reinforcing plate. Spacing between two splices or between a splice and a reinforcing plate should be no less than three times the length of the longer splice. Splicing under minor fittings such as drag wire, antidrag wire, or compression strut fittings is acceptable under the following conditions:
a. The reinforcement plates of the splice should not interfere with the proper attachment or alignment of the fittings. Do not alter the locations of pulley support brackets, bellcrank support brackets, or control surface support brackets. Plates are to be tapered off, as depicted in Figure 1-2.
b. The reinforcement plate may overlap drag wire, antidrag wire, or compression strut fittings, if the reinforcement plates are on the rear face of the rear spar or the front face of the front spar. In such cases, it will be necessary to install slightly longer bolts. The front face
reinforcement plate should not overlap drag strut fittings, except when it does not require sufficient shortening of compression struts or changes in drag-truss geometry, to prevent adjustment for proper rigging. Even though take up is sufficient, it may be necessary to change the angles on the fittings. (Acceptable methods for splicing the various types of spars are shown in Figure 1-4 through Figure 1-9.) Reinforcement plates must be used as indicated on all scarf repairs to spars and the slopes of scarves shown are minimum slopes.
FIGURE 1-5. METHOD OF SPLICING SOLID OR LAMINATED
RECTANGULAR SPARS → 6A RECOMMENDED * + 2A+ 5A MINIMUM 6A RECOMMENDED
• 15A MINIMUM → SA MINIMUM
NO FITTINGS WITHIN THESE LIMITS 1/4 A
DIRECTION OF GRAIN IF SPRUCE
OR OUTER FACE GRAIN IF PLYWOOD Reinforcement plates to be spruce or plywood and should be bonded only, with the ends feathered off at 5:1 slope as depicted in Figure 1-2.
Solid spars may be replaced with laminated spars or vice versa, provided the material is of the same high quality.
Par 1-40 Page 1-18
FIGURE 1-6. METHOD OF SPLICING SOLID T" SPARS If splice is made where routing is feathered to full width of spar, tapered plates (conforming to the contour of the routing) should be added first;
otherwise, the splice is the same as shown.
6A • 6A 6A 6A 15A MINIMUM SPLICE
NO FITTING WITHIN THESE LIMITS B ½B
DIRECTION OF GRAIN Reinforcement plates may be of plywood, the same material as the spar, or material of equal or higher quality; and should be bonded only with ends feathered out with a 5:1 slope.
Par 1-40 Page 1-19
FIGURE 1-7. REPAIRS TO BUILT-UP "T" SPAR
- 15A 24 2A
NO FITTING WITHIN THESE LIMITS DIRECTION OF GRAIN IN ¼B
PLYWOOD REINFORCEMENT
PLATES TO BE SAME AS ORIGINAL WEB.
PLYWOOD
A - SOLID WOOD FILLER BLOCKS Ends of reinforcement plates should be feathered out Section A- with a 5:1 slope.
Par 1-40 Page 1-20
FIGURE 1-8. METHOD OF SPLICING BOX SPAR FLANGES (PLATE METHOD)
• NEW WEB ANGLE OF FACE GRAIN SAME AS IN ORIGINAL WEB.
STAGGER WEB SPLICES. (See figure 1-9 for method of splicing plywood webs.)
15B D
FOR CLARITY
• 15A - 6A NEW WEB NOT
SPRUCE REINFORCEMENT PLATES INCLUDED IN THIS VIEW A, B, C, D = ORIGINAL DIMENSIONS Par 1-40 Page 1-21
FIGURE 1-9. METHOD OF SPLICING BOX SPAR WEBS
- FILLER BLOCK + 10A→ - COVER STRIP *10A+1* • INSERTED WEB SECTION
COVER STRIPS Same Thickness as "A" COVER STRIPS
ORIGINAL SPAR WEB
INSERTED WEB
SECTION FILLER BLOCK→ - FILLER BLOCK
1. AFTER INSERTED WEB HAS BEEN BONDED IN PLACE, BOND COVER STRIPS OVER ENTIRE LENGTH OF WEB SPLICE JOINTS.
2. SECTIONAL SHAPE OF FILLER BLOCKS MUST CONFORM EXACTLY TO TAPER OF SPAR. THEY MUST
NOT BE TOO TIGHTLY FITTED, FOR WEDGING ACTION WILL LOOSEN EXISTING BONDS OF WEBS TO FLANGES. IF TOO LOOSELY FITTED, CRUSHING OF WEB WILL OCCUR WHEN CLAMPING.
3. DIRECTION OF FACE GRAIN OF INSERTED WEB SECTION AND COVER STRIPS TO BE SAME AS ORIGINAL WEB.
1-41. SPAR REPLACEMENT. Replacement of spars is a major repair. Spars may be replaced by new parts made by the manufacturer or the holder of a Parts Manufacturer Approval (PMA) for that part. Owner-produced spars may be installed providing they are made from a manufacturer-approved drawing. Also, a spar may be made by reference to an existing spar providing sufficient evidence is presented to verify that the existing spar is an original part, and that all materials and dimensions can be determined. The dimensions and type of wood used are Par 1-40 Page 1-22
Par 1-41 Page 1-23 critical to the structural strength of the aircraft. Care should be taken that any replacement spars accurately match the manufacturer’s original design.
1-42. SPLICING OF BOX SPAR WEBS. Always splice and reinforce plywood webs with the same type of plywood as found on the original part. Do not use solid wood to replace plywood webs. Plywood is stronger in shear than solid wood of the same thickness due to the grain direction of the individual plies. The face-grain of plywood replacement webs and reinforcement plates must be in the same direction as the original member to ensure that the new web will have the required strength. (The method of splicing plywood webs is shown in Figure 1-9.)
1-43. REPLACING SOLID-TYPE SPARS WITH LAMINATED-TYPE SPARS. Solid spars may be replaced with laminated spars or vice versa, provided the material is of the same high quality. External reinforcements (plywood or solid) must always be replaced as on the original member.
1-44. SPAR LONGITUDINAL CRACKS AND LOCAL DAMAGE. Cracked spars (except box spars) may be repaired by bonding plates of spruce or plywood of sufficient thickness to develop the longitudinal shear on both sides of the spar. Extend the plates well beyond the termination of the cracks, as shown in Figure 1-10. A method of repairing small local damage to either the top or bottom side of a spar is also shown in Figure 1-10.
Par 1-44 Page 1-24 FIGURE 1-10. METHOD OF REINFORCING A LONGITUDINAL CRACK AND/OR
LOCAL DAMAGE IN A SOLID SPAR
a.
Longitudinal Cracking of Wood Wing Spars of Aircraft Operating in Arid Regions. Aircraft having wood spars and operating in arid regions may develop longitudinal spar cracks in the vicinity of the plywood reinforcement plates. These cracks result from the tendency of the spar to shrink when drying takes place. Plywood resists this tendency to shrink and causes the basic spar stock to split (see paragraph 1-2c). Cracks start under the plywood plates, usually (but not necessarily) at a bolt hole or cutout, and usually spread in each direction until they extend a short distance beyond the ends of the plates where the resistance to spar shrinkage disappears. Cracks have also been found in the butt end of spars. Other factors, which have been found conducive to the formation of cracks are poor protective finishes, large cutouts, and metal fittings that utilize two lines of large diameter bolts.
b. Repairing Cracks Versus Installing a New Spar. The presence of cracks does not necessarily mean that the spar must be discarded. If the crack is not too long or too close to either edge and can be reinforced properly, it will probably be more economical and satisfactory to perform repair rather than install a new spar or section. However, a generally acceptable procedure suitable for all airplane models is not available. Because of the possibility of strength deficiencies contact the manufacturer. In absence of the manufacturer, the FAA should be