Par 2-30 Page 2-67
SECTION 3. INSPECTION AND TESTING
GENERAL. All components of the covering should be inspected for general condition.
Loose finishing tape and reinforcing patches;
chafing under fairings; brittle, cracking, peeling, or deteriorated coatings; fabric tears and rock damage; broken or missing rib lacing;
and rodent nests are unacceptable. The entire fabric covering should be uniformly taut with no loose or wrinkled areas, or excess tension which can warp and damage the airframe.
Excess Tension. There are no methods or specifications for measuri ng acceptable fabric tension other than observation. Excess tension may pull and warp critical components, such as: longerons; wing ribs;
and trailing edges out of position, weakening the structure.
Excess tension with cotton, linen, and glass fiber fabric covering is usually caused by excessive dope film on a new covering, or continuous shrinking of an originally satisfactory dope film as the plasticizers migrate from the dope with age. Heat from sun exposure accelerates plasticizer migration.
Excess tension with polyester fabric, coated with dope, is usually caused by the combined tension of the heat tautened polyester fabric and continuous shrinking of the dope film as the plasticizers migrate from the dope with age.
Loose Fabric. Fabric that flutters or ripples in the propeller slipstream, balloons, or is depressed excessively in flight from the static position, is unacceptable.
Loose or wrinkled cotton, linen, and glass fabric covering may be caused by inadequate dope film; poor quality dope; fabric installed with excess slack; or by a bent, broken, or warped structure.
Loose or wrinkled polyester fabric covering, finished with coatings other than dope, may be caused by inadequate or excessive heat application; excess slack when the fabric was installed; or bent or warped structure.
Polyester fabric which does not meet aircraft quality specifications will very likely become loose after a short period of time.
Glass fabric covering should be tested with a large suction cup for rib lacing cord failure and reinforcing tape failure caused by chafing on all wing ribs and other structural attachments throughout the airframe. Particular attention should be given to the area within the propeller slipstream. If failure is indicated by the covering lifting from the static position, the rib lacing cord and reinforcing tape must be reinstalled with double the number of original laces.
NOTE: Temporary wrinkles will develop in any fabric coated and finished with dope, when moisture from rain, heavy fog, or dew is absorbed into a poor-quality dope film, causing the film to expand.
Temporary wrinkles may also develop with any type of thick coatings, on any type of fabric, when an aircraft is moved from a cold storage area to a warm hangar or parked in the warming sunshine, causing rapid thermal expansion of the coating.
Coating Cracks. Fabric exposed through cracks in the coating may be initially tested for deterioration by pressing firmly with a thumb to check the fabric’s strength. Natural fibers deteriorate by exposure to ultraviolet radiation, mildew, fungus from moisture, high acid-content rain, dew, fog, pollution, and age.
Polyester filaments will deteriorate by exposure to UV radiation.
Par 2-32 Page 2-68 Glass fabric will not deteriorate from UV exposure, but will be deteriorated by acid rain, dew fallout, and chaffing if loose in the prop blast area.
Cotton, linen, and glass fabric coverings are dependent solely on the strength and tautening characteristics of the dope film to carry the airloads. Dope coatings on heattautened polyester fabric will also absorb all the airloads because the elongation of polyester filaments are considerably higher than the dope film. Polyester fabric that is coated with materials other than dope, is dependent solely on the heat tautening and low-elongation characteristics of the polyester filaments to develop tension and transmit the airloads to the airframe without excess distortion from a static position.
Cracks in coatings will allow any type of exposed fabric to deteriorate. Cracks should be closed by sealing or removing the coatings in the immediate area and replace with new coatings, or recover the component.
FABRIC IDENTIFICATION.
Cotton Fabric meeting TSO-C15 or TSO-C14 can be identified by an off-white color and thread count of 80 to 94 for TSO-C14b and 80 to 84 for TSO-C15d in both directions.
Aircraft linen conforming to British specification 7F1 may be identified by a slightly darker shade than cotton fabric and irregular thread spacing. The average thread count will be about the same as Grade A fabric (TSO-C15d). The non-uniformity of the linen thread size is also noticeable, with one thread half the size of the adjacent thread. When viewed under a magnifying glass, the ends of the cotton and linen fiber nap may be seen on the backside. The na p is also seen when the coating is removed from the fr ont or outside surface. A light-purple color showing on the back side of cotton or linen fabric indicates a fungicide was present in the dope to resist deterioration by fungus and mildew.
Polyester fabric conforming to TSO-C14b or TSO-C15d is whiter in color than cotton or linen. The fabric styles adapted for use as aircraft covering have a variety of thread counts, up to ninety-four (94), depending on the manufacturing source, weight, and breaking strength. Polyester is a monofilament and will not have any nap or filament ends showing.
Glass fabric is manufactured white in color, and one source is precoated with a blue-tinted dope as a primer and to reduce weave distortion during handling. Thread count will be approximately 36 threads per inch. Glass fabrics are monofilament and will not have any nap or filament ends showing unless they are inadvertently broken.
When a small fabric sample can be removed from the aircraft and all the coatings removed, a burn test will readily distinguish between natural fabric, polyester, and glass fabric. Cotton and linen will burn to a dry ash, polyester filaments will melt to a liquid and continue burning to a charred ash, and glass filaments, which do not support combustion, will become incandescent over a flame.
COATING IDENTIFICATION.
Nitrate or butyrate dope must be used to develop tension on cotton, linen, and glass fabrics. When a small sample can be removed, burn tests will distinguish nitrate dope-coated fabric from butyrate dope-coated fabric by its immediat e ignition and accelerated combustion. Butyrate dope will burn at less than one-half the rate of nitrate dope. Coating types other than nitrate or butyrate dope may have been used as a finish over dope on cotton, linen, and glass fiber fabric coverings.
Par 2-32 Page 2-69 If the fabric type is determined to be polyester, coating identification should start by reviewing the aircraft records and inspecting the inside of the wings and the fuselage for the required fabric source identification stamps for covering materials authorized under the STC. The manual, furnished by the holder of the STC-approved fabric, should be reviewed to determine whether the coati ngs are those specified by the STC.
Coating types, other than those authorized by the original STC, may have been used with prior FAA approval, and this would be noted in the aircraft records. The presence of dope on polyester can be detected by a sample burn test.
STRENGTH CRITERIA FOR AIRCRAFT FABRIC.
Minimum performance standards for new intermediate-grade fabric is specified in TSO-C14b, which references AMS 3804C.
Minimum performance standards for new Grade A fabric are specified in TSO-C15d, which references AMS 3806D.
The condition of the fabric covering must be determined on every 100-hour and annual inspection, because the strength of the fabr ic is a definite factor in the airworthiness of an airplane. Fabric is considered to be airworthy until it deteriorates to a breaking strength less than 70 percent of the strength of new fabric required for the aircraft. For example, if grade-A cotton is used on an airplane that requires only intermediate fabric, it can deteriorate to 46 pounds per inch width (70 percent of the strength of intermediate fabric) before it must be replaced.
Fabric installed on aircraft with a wing loading less than 9 lb. per square foot (psf), and a Vne less than 160 mph, will be considered unairworthy when the breaking strength has deteriorated below 46 lb. per inch width, regardless of the fabric grade. Fabric installed on aircraft with a wing loading of 9 lb. per square foot and over, or a Vne of 160 mph and over, will be considered unairworthy when the breaking strength has deteriorated below 56 lb. per inch width.
Fabric installed on a glider or sailplane with a wing loading of 8 lb. per square foot and less, and a Vne of 135 mph or less, will be considered unairworthy when the fabric breaking strength has deteriorated below 35 lb. per inch width, regardless of the fabric grade.
FABRIC TESTING. Mechanical devices used to test fabric by pressing against or piercing the finished fabric are not FAA approved and are used at the discretion of the mechanic to base an opinion on the general fabric condition.
Punch test accuracy will depend on the individual device calibration, total coating thickness, brittleness, and types of coatings and fabric. Mechanical devices are not applicable to glass fiber fabric that will easily shear and indicate a very low reading regardless of the true breaking strength. If the fabric tests in the lower breaking strength range with the mechanical punch tester or if the overall fabric cover conditions are poor, then more accurate field tests may be made.
Cut a 1-1/4-inch wide by 4-inch long sample from a top exposed surface, remove all coatings and ravel the edges to a 1-inch width.
Clamp each end between suitable clamps with one clamp anchored to a support structure while a load is applied (see table 2-1) by adding sand in a suitable container suspended a few inches above the floor. If the breaking strength is still in question, a sample should be sent to a qualified testing laboratory and breaking strength tests made in accordance with American Society of Testing Materials (ASTM) publication D5035.
Par 2-42 Page 2-70 NOTE: ASTM publication D1682 has been discon tinued but is still referred to in some Aerospace Material Specification (AMS). Th e grab test method previously listed in ASTM D1682, sections 1 through 16, has been superseded by ASTM publication D5034.
The strip testing method (most commonly used in aircraft) previously listed in ASTM D1682, sections 17 through 21, has been superseded by ASTM publication D5035.
REJUVENATION OF THE DOPE FILM. If fabric loses its strength, there is nothing to do but remove it and recover the aircraft. But if the fabric is good and the dope
Par 2-42 Page 2-71 is cracked, it may be treated with rejuvenator, a mixture of very potent solvents and plasticizers, to restore its resilience. The surface of the fabric is cleaned and the rejuvenator sprayed on in a wet coat, and the solvents soften the old finish so the plasticizers can become part of the film. When the rejuvenator dries, the surface should be sprayed with two coats of aluminum-pigmented dope, then sanded and a third coat of aluminum-pigmented dope applied, followed with the colored-dope finish.
When repairing, rejuvenating, and refinishing covering materials approved under an STC, instructions in the manual furnished by the material supplier should be followed.
2-36.2-41. [RESERVED.]
Par 2-42 Page 2-72 SECTION 4. REPAIRS TO FABRIC COVERING
GENERAL. All materials used to make repairs to fabric covering must be of a quality at least equal to the original materials. Workmanship and repair methods must be made in a manner that will return the fabric covering to its original airworthy condition.
Any combination of seams hand-sewn and overlapped and doped may be used to ma ke repairs and install new fabric sections. (See paragraph 2-7.)
All pigmented dope coats, including aluminumpigmented coats, should be removed to the clear dope preliminary coats before installing a new fabric section or finishing tape on the old fabric. The appropriate dope thinner (see paragraph 2-20) may be applied with a brush to soften the old dope. The softened coats can then be removed down to the clear dope coats by scraping with a dull-bladed knife while supporting the fabric from the back side. Removing the old dope by sanding is optional.
Avoid allowing dope to run down the back side of the fabric or drip through the wing onto the back side of the opposite surface, which will cause cosmetic damage and will show in a high gloss finish.
Repairing a new fabric section over two adjacent wing ribs is considered a major repair. A log book entry and an FAA Form 337 must be processed accordingly.
All fabric patch edges not covered with a finishing tape should have a pinked e dge or a 1/4-inch raveled edge.
Where the edge of a new fabric section will be located within 1 inch of a structural member to which the fabric is attached by rib lacing or other methods, the new fabric section should be extended 3 inches past the structural member.
NOTE: Before installing new finishing tape, duplicate the original rib lacing or other attachments without removing the original rib lacing or attachment. Removing the original finishing tape is optional.
When installing large sections of fabric on a wing or other components, all machine-sewn and/or overlapped and doped seams should be made in accordance with the guidelines specified in paragraphs 2-7 and 2-8.
When repairing a covering material other than cotton or linen, which was approved with the manufacturer’s type certificate (TC), or approved under the authority of an STC, follow the repair instructions furnished by the aircraft manufacturer or supplier of the STC-approved covering materials.
REPAIR OF TEARS AND ACCESS OPENINGS.
When all the original fabric is intact, an opening may be repaired by sewing the two sides together with a curved needle as illustrated in figures 2-13 and 2-14. The fabric edges should be pulled together uniformly with no wrinkles. Before sewing, remove the old dope coats down to the clear dope coats a distance of 2 inches on each side of the opening. The hand-sewn thread quality should be at least equal to that specified in table 2-2 and treated with wax (paraffin-free or beeswax) to prevent fraying, or use the proper thread on the STC application. After sewing, apply a coat of clear dope over the cleaned area and install a 3-inch wide finishing tape, centered over the stitches.
Par 2-42 Page 2-73
FIGURE 2-13. Repair of tears in fabric.
Finishing tapes should be well saturated with dope and smoothed out with no voids or wrinkles during installation. After drying for 1 hour at 70 °F, additional coats of clear dope followed with pigmented dope are applied as detailed in paragraph 2-21.
If the opening is more than 8 inches long but less than 16 inches long in any direction, the finishing tape width should be increased to 4 inches.
The finishing tape width should be increased to 6 inches if the opening is more than 16 inches long in any direction, is located on a wing top surface, or the aircraft Vne speed is greater than 150 mph.
SEWN-PATCH REPAIR. Openings that cannot be repaired by closing with stitches may be repaired by sewing in a new fabric section.
The edges of the fabric around the opening should be trimmed straight on four sides to facilitate the installation of straight sections of finishing tape over the stitches.
After cutting out the damaged section and removing the coatings as detailed in paragraph 2-42, the new fabric section should be sized to allow folding both edges of the fabric back 1/2-inch to increase the stitch tear resistance. Temporarily attach the four corners in position with thread. Start with a double thread with a square knot at the end (see figure 2-14) and continue stitching in the manner described in figures 2-13 and 2-14. When the stitching is complete, wet the new fabric section as described in paragraph 2-21 to remove any creases. After drying, apply one coat of clear dope on the new fabric, as described in paragraph 2-21.
When the first coat of dope has dried 1 hour at 70 °F, apply a 3-inch wide finishing tape, centered over the stitches. The finishing ta pe should be well saturated with
Par 2-44 Page 2-74 Step 1. Tie the thread ends together with a square knot.
Step 2. Start the stitch by routing through the loop, positioning the square knot on the inside.
Step 3. Complete the hand-stitch and secure the stitching with a half hitch at a maximum each 10 stitches.
FIGURE 2-14. Hand-stitch detail.
dope and smoothed out with no voids or wrinkles.
Additional coats of clear dope and pigmented dope are then applied to obtain the desired tautness and finish, as described in paragraph 2-21.
If the opening is more than 8 inches but less than 16 inches long in any direction, the finishing tape should be 4 inches wide.
The finishing tape width should be 6 inches wide if the opening is over 16 inches long in any direction, is located on a wing top surface, or the aircraft Vne speed is greater than 150 mph.
DOPED-ON PATCH REPAIR. An opening not over 8 inches in length in any direction, on an aircraft with a Vne speed less than 150 mph, may be repaired with a 2-inch overlapped and doped patch. The opening should be trimmed to eliminate any irregular edges and old pigmented dope coats removed as described in paragraph 2-42.
When installing a fabric patch over a small opening, the loose edge of the fabric around the opening may be secured by extending a series of small threads, from the edge across the opening, to the opposite side. After the patch is completed, the threads may be clipped and removed through an access port or left in place.
The fabric patch is installed by applying a coat of clear dope around the opening, then positioning the patch over the opening. Brush out any void or wrinkles while saturating only the fabric overlap area. After the first coat of clear dope around the edge has dried 1 hour at 70 °F, wet the fabric patch to remove any creases as described in paragraph 2-21. After drying, apply additional coats of clear dope and pigmented dope over the entire patch as described in paragraph 2-21.
If the opening is less than 8 inches in length in any direction and the aircraft Vne speed is greater than 150 mph, a 2-inch wide finishing tape should be installed on all sides, centered on the edge of the 2-inch overlap patch.
If the opening is more than 8 inches but less than 16 inches in length, in any direction on an aircraft with a Vne speed less than
Par 3-1 Page 3-75 150 mph, it may be repaired with a doped patch, which is overlapped 1/4 of the opening maximum dimension. The maximum overlap should not exceed 4 inches.
If the opening is more than 8 inches but less than 16 inches in length in any direction, the repair is located on a wing top surface, and the aircraft Vne speed is greater than 150 mph, the patch overlap should be 4 inches and a 2-inch wide finishing tape installed on all sides, centered on the edge of the patch.
Par 3-1 Page 3-76 If the opening is more than 16 inches in length in any direction and the Vne speed is less than 150 mph, the patch overlap should be 4 inches and the finishing tape should be 4 inches in width, centered on the edge of the patch.
If the opening is more than 16 inches in length in any direction and the Vne speed is greater than 150 mph, the patch overlap should be 4 inches and the finishing tape should be 6 inches in width, centered on the edge of the patch.
2-46.2-51. [RESERVED.]
Par 3-1 Page 3-77 CHAPTER 3. FIBERGLASS AND PLASTICS
SECTION 1. REPAIR OF LIGHT LOAD LAMINATE STRUCTURES
GENERAL. There is a wide variation in the composition and structural application of laminates, and it is essential that these factors be given major consideration when any restoration activities are undertaken. To a similar extent, there also exist many types of laminate structure repairs that may or may not be suitable for a given condition. For this reason, it is important that the aircraft or component manufacturer’s repair data be reviewed when determining what specific type of repair is permissible and appropriate for the damage at hand.
NOTE: Review Material Safety Data Sheets for material to be used. When handling materials, prepreg fabrics, or parts with prepared surfaces, observe shelf life. Latex gloves and approved masks must be worn.
The materials used in the repair of laminate structures must preserve the strength, weight, aerodynamic characteristics, or electrical properties of the original part or assembly. Preserva tion is best accomplished by replacing damaged material with material of identical chemical composition or a substitute approved by the manufacturer.
To eliminate dangerous stress concentrations, avoid abrupt changes in cross-sectional areas. When possible, for scarf joints and facings, make small patches round or oval-shaped, and round the corners of large repairs.
Smooth and properly contour aerodynamic surfaces.
Test specimens should be prepared during the actual repair. These can then be s ubjected to a destructive test to establish the quality of the adhesive bond in the repaired part. To make this determination valid, the specimens must be assembled with the same adhesive batch mixture and subjected to curing pressure, temperature, and time identical to those in the actual repair.
FIBERGLASS LAMINATE REPAIRS. The following repairs are applicable to fiberglass laminate used for non-structural fairing, covers, cowlings, honeycomb panel facings, etc. Prior to undertaking the repair, remove any paint by using normal dry sanding methods. Bead blasting may be used but caution must be exercised not to abrade the surfaces excessively.
NOTE: Chemical paint strippers must not be used.
NOTE: These repairs are not to be used on radomes or advanced composite components, such as graphite (carbon fiber) or Kevlar.
CAUTION: Sanding fiberglass laminates gives off a fine dust that may cause skin and/or respiratory irritation unless suitable skin and respiration protection is used.
Sanding also creates static charges that attract dirt or other contaminants.
Check for voids and delamination by tap testing. (See chapter 5.) When the surface of a fiberglass laminated structure is scratched, pitted, or eroded; first wash with detergent and water to remove all of the dirt, wax, or oxide film. Then scrub th e surface with an acceptable cleaner. After the surface is thoroughly cleaned, sand it with 280-grit sandpaper, and again use an acceptable cleanser to remove any sanding residue and moisture.
This is
Par 3-3 Page 3-78 essential, as any moisture remaining on the surface will inhibit the cure of the resin. Dry the fiberglass laminate thoroughly prior to bonding repair. Mix enough resin, using the manufacturer’s instructions, to completely cover the damaged area, and apply one or two coats.
Cover the resin with a peel ply to exclude all air from the resin while it is curing. After the resin has cured, remove the film and file or sand the surface to conform to the original shape of the part. Ensure that all edges of the laminate part are sealed to prevent water absorption. Then refinish it to match the rest of the structure.
Superficial scars, scratches, surface abrasion, or rain erosion can generally be re paired by applying one or more coats of a suitable low temperature resin, catalyzed to cure at room temperature, to the abraded surface. The number of coats required will depend upon the type of resin and the severity of the damage.
Damage not exceeding the first layer or ply of fiberglass laminate can be repaired by filling with a putty consisting of a compatible roomtemperature-setting resin and clean short glass fibers. Before the resin sets, apply a sheet of peel ply over the repair area and work out any bubbles and excess resin. After the resin has cured, sand off any excess and prepare the area for refinishing.
Damage deep enough to seriously affect the strength of the laminate (usually more than the first ply or layer of fabric) may be repaired as illustrated in figure 3-1. Coat the sanded area with room-temperature-setting resin and apply contoured pieces of glass fabric soaked in resin. Apply a peel ply sheet over the repair and work out any bubbles and excess resin.
After the resin has cured, scrape off the excess resin and sand the surface of the repair to the original contour.
FIGURE 3-1. Typical laminate (facing) repair.
Damage that extends completely through one facing and into the core requires the replacement of the damaged core and facing.
A method for accomplishing this type of repair is shown in figure 3-2. An alternate method for repairing the facing is shown in figure 3-3.
The damaged portion is carefully trimmed out to a circular or oval shape and the core material is removed completely to the opposite facing. Exercise caution so as not to damage the opposite facing or to start delamination between the facings and the core around the damage.
FIGURE 3-2. Typical core and facing repair.