slowly until a new equilibrium is reached, as shown Room temperature equilibrium in figure 5-99.
Viscous cement Viscous cement applied.
Joint dried at room temp., still contains solvent.
Solvent starts to form Equilibrium after cushions,. heat treatment Under pressure Bead extruded, Bead cleaned away Joint "drying" Heat treatment expands cushion, reduces Cushions hardening, concentration of solvent in joint.
FIGURE 5-98. Cementing with syrup cement. FIGURE 5-99. Effect of heat treatment.
On cooling, the cushion will be larger and correspondingly harder since it contains less solvent per unit of volume. Heating a solvent joint long enough to expand its cushion, therefore, will produce a much stronger joint.
Cemented joints must be heat treated with caution. Heat first activates the solvent, which softens the cushion. The cushion then slowly expands as the solvent penetrates the material. In heat treating, it is important that the temperature does not approach the softening point of transparent plastics.
It is important that the joint be thoroughly hardened before machining, sanding, or polishing to remove the bead.
Care and Maintenance Vision is so vital in aircraft that day-to-day maintenance of transparent enclosures is of the utmost importance. Proper maintenance methods should be carried out thoroughly whenever vision is impaired as a result of chemical or physical actions or defects, and every effort should be put forth to eliminate harmful action while servicing the aircraft.
The replacement of transparent plastic enclosures has been necessitated by severe crazing, apparently caused by exposure to harmful solvents and improper maintenance handling. The crazing appears as a network of cracks running in all directions over the surface of the plastic. It can also occur within the plastic at or near cemented joints.
The use of improper cleaning fluids or compounds is one of the most common causes of these dificulties. The crazing action of a solvent is often delayed; that is, crazing may not appear for several weeks after the exposure to solvent or fumes. It is not always possible to determine immediately, by simple trial, whether a particular cleaner will be injurious or not. To minimize the damage, the precautions discussed in the following paragraphs should be observed.
Routine removal of film and other operational soils, where abrasive polishing for scratch removal is not required, can be accomplished by the use of aqueous detergent solutions. Two recommended solutions are: wetting agent, synthetic, nonionic, conforming to Military Specification MIL-D-16791; or wetting agent, alkyl aryl sulfonate, 40% active.
These materials should be used in concentrations of 2 or 3 oz. per gal, of water. They should be applied 218
with soft cloths or photographic cellulose sponges which have been used for no other purpose. Polish and cleaner conforming to Military Specification MIL-C-18767 will give satisfactory results for most cleaning requirements.
When cleaning exterior surfaces, always remove rings from the hands before washing the transparent plastic. The cleaning procedure is comprised of the following steps:
that the platsic can become considerably hotter than the surrounding air inside or outside the aircraft.
Air inside an unshaded and unventilated aircraft will transfer the heat radiated by the metal members in the aircraft to the plastic enclosure by convection.
To prevent heat deformation of transparent plastic enclosures on aircraft parked exposed to the sun, the following precautions are recommended:
(1) Flush the plastic surface with plenty of (1) If surrounding air temperature is below water, using the bare hands to feel for 100° F., no special precautions are necesand gently dislodge any dirt, sand, or sary.
mud. (2) If surrounding air temperature is between
(2) Wash with mild soap and water. Be sure 100° and 120° F., enclosures should be the water is free of harmful abrasives. A opened sufficiently to permit free circulasoft cloth, sponge, or chamois may be tion of air through the aircraft and under used in washing, but only to carry the the enclosure.
soapy water to the plastic. Go over the (3) If the surrounding air temperature is surface with bare hands to quickly detect above 120° F., the enclosure must be and remove any remaining dirt before it opened and protected from the sun by a scratches the plastic. suitable cover which does not come into
(3) Dry clean contact with the transparent plastic. If soft oni tin ue possible, the aircraft should be parked in rubbin ter it the shade.
dry. ut may (4) To remove enclosure covers, lift them off;
build which sliding may cause abrasion of the plastic attract face be- surfaces.
comes with Compounds for paint stripping, degreasing, and the ch brightening, as well as most organic solvents, cause
(4) Never serious damage to transparent acrylic plastics. All polishi such parts should be removed before starting paint The procedur stripping, and should not be replaced until the sists of three steps: cleaning and painting is completed and the paint or lacquer is thoroughly dry, since paint and lacquer cause crazing of plastics. The plastic parts should be removed from the area where the stripping, degreasing, or painting is being done. The parts should be protected with soft cloth covers.
If it is impracticable to remove a plastic panel, cut a polyethylene sheet (minimum of 0.010-in.
thick and containing no pinholes) to match as exactly as possible the size of the window. The polypark ethylene sheet should fit snugly over the surface of and the plastic window, and the edges should be caretake fully taped with masking tape at least 2 in. wide to park permit at least 1 in. of sealing width on both the three sources. plastic film and the aircraft. Make certain that no property of liquid or fumes can seep through to the window. It selec lucing rays of is important that the entire surface of the window 219
be covered and that no cutting tools be used to remove the masking.
Aluminum foil is unsatisfactory as a protection from paint (and other sprays which contain solvents) because of its low resistance to tears, punctures, and pinholes. Protective coating conforming to Military Specification MIL-C-6799 is satisfactory as a protection from paint and other sprays which contain solvents.
Do not sand transparent plastics unless it is absolutely necessary. Hairline scratches of 0.001-in.
maximum depth should be left as is, provided optical requirements are maintained.
INSTALLATION PROCEDURES There are a number of methods for intalling transparent plastic panels in aircraft. The method the aircraft manufacturer utilizes depends on the position of the panel on the aircraft, the stresses to which it will be subjected, and a number of other factors. When installing a replacement panel, follow the same mounting method used by the aircraft manufacturer.
Where difficulty is encountered in rivet installation, bolts may be substituted when installing replacement panels, provided the manufacturer's original strength requirements are met and the bolts do not interfere with adjoining equipment.
In some instances replacement panels do not fit the installation exactly. Whenever adjustment of a replacement panel is necessary, the original design drawing, if available, should be consulted for proper clearances. The following principles should be considered in installing all replacement panels.
Fitting and handling should be done with masking material in place. Do not scribe plastic through masking material. On edges where transparent materials will be covered, or attached to, remove the masking material. When subject to large stresses, transparent plastics are apt to craze. It is of prime importance that plastics be mounted and installed so that such stresses are avoided.
Since transparent plastic is brittle at low tempera tures, extra care must be taken to prevent cracking during maintenance operations. Transparent plastic parts should be installed at room temperature, if Never force a transparent panel out of shape to make it fit a frame. If a replacement does not fit easily into the mounting, obtain a new replacement or sand the panel sufficiently to obtain the exact size that conforms with the mounting frame.
Do not heat and re-form areas of the panel, since local heating methods are likely to be only superficial and not thorough enough to reduce stress concentrations.
Since plastics expand and contract approximately three times as much as metal, suitable allowance for dimensional changes with temperature must be made. Use the values shown in figure 5-100 as minimum clearances between the frames and the plastics.
Dimensional Allowance in Inches * Required for Expansion Required for Contraction Dimension of Panel from 25°C (77°F) to from 25°C (77°F) to in Inches ** 70°C (158°F) -55°C (-67°F) 12 0.031 0.050 24 0.062 0.100 36 0.093 0.150 48 0.124 0.200
0.155 0.250 72 0.186 0.300
• Where the configuration of a curved part is such as to take up dimensional changes by change of contour, the allowances given may be reduced if it will not result in localized stress. Installations permitting linear change at both ends require half the listed clearances.
** For dimensions other than those given use proportional clearance.
FIGURE 5-100. Expansion and contraction allowances.
Bolt and Rivet Mountings In bolt installations, spacers, colars, shoulders, holders or stop nuts should be used to prevent excessive tightening of the bolt. Whenever such devices are used by the aircraft manufacturer, they should be retained in the replacement installations.
To ensure long service, give special consideration to the following factors:
(1) Use as many bolts or rivets as practical.
(2) Distribute the total stresses as equally as possible along the bolts and rivets.
(3) Make sure the holes drilled in the plastic are sufficiently larger than the diameter of the bolt to permit expansion and contraction of the plastic relative to the frame.
(4) Make sure the holes in the plastic are concentric with the holes in the frame so that the greater relative expansion of the plastic will not cause binding at one edge of the hole. The hole should be smooth and free of any nicks or roughness.
(5) Use oversize tube spacers, shoulder bolts, rivets, cap nuts, or some other device to protect the plastic from direct pressure.
Synthetic Fiber Eage Attachment Modern edge attachments to transparent plastic assemblies are made of synthetic fibers specially impregnated with plastic resins. The most commonly used fibers are glass, orlon, nylon, and dacron.
Reinforced laminated edge attachments are the preferred type, especially when mounting by bolts or rivets is necessary. The edges have the advantage of more efficiently distributing the load and reducing failures caused by differential thermal expansion.
Laminated edge attachments can be mounted by any of the foregoing methods, with any needed holes drilled through the edge attachment material and not the transparent plastic.
The most efficient method of mounting a laminated edge attachment is by the "slotted hole" method. The slotted holes are in the edge attachment and allow for differential thermal expansion.
Fabric loop attachments are sometimes attached to the plastic material with a cable or extrusion contained within the loop. A special extrusion is necessary to contain the loop and cable.
LAMINATED PLASTICS Laminated plastic enclosures are made by bonding two layers of monolithic transparent sheets together with a soft plastic inner layer. They are installed in pressurized aircraft because of their superior shatter resistance and greater resistance to explosive decompression as compared to monolithic plastic enclosures.
Cellulose Acetate Base Plastics In general, the methods used for fabrication, repair, and maintenance of cellulose acetate base plastics are similar to those used for acrylic plastics. In handling cellulose acetate base plastics, give attention to the following variations and additions to the recommendations already given for acrylic plastics.
Since the chemical composition of acetate base plastics differs greatly from that of acrylics, the cement used is of a different type. Generally, two types are used, solvent and dope.
Solvent type cement is generally used where transparency must be maintained in the joint. It is relatively quick drying and is well adapted for use in making emergency repairs. However, even though the cement is quick drying, the drying time will vary with the size of the joint and atmospheric conditions. Acetone may be used as a solvent type cement.
Dope type cement is preferred for use where the surfaces to be joined do not conform exactly. This cement softens the surfaces of a joint and, at the same time, creates a layer between the two pieces being cemented. However, it does not give a transparent joint and is slower drying than the solvent cement. It will take from 12 to 24 hrs. for the joint to reach full strength.
Since the expansion and contraction rates of acetate-base plastics are greater than those of acrylics, make greater allowances when mounting them.
These plastics are affected by moisture and will swell as they absorb water. In general, allow ⅛ in.
per foot of panel length for expansion, and ¾e in.
per foot for contraction.
FIBER GLASS COMPONENTS Because of the unequaled strength/weight ratio, the ability to pass radio and radar waves, the ease of manufacture in contoured shapes, immunity to mildew and weather resistant characteristics, and adaptability to numerous places and shapes, fiber glass is a versatile material with numerous uses in modern aircraft construction. A few of the many applications are radomes, radio antenna covers, and junction boxes.
Fiber glass is manufactured from specially processed glass balls. By a fabrication process, the glass is turned into fibers which may result in an end 221
product of cloth, molded mats, or yarn. The yarn is used to manufacture molded parts. The fiber glass cloth is used in making laminated shapes or in the repair of laminated assemblies. Another use is in the repair of metal structures.
Mat Moided Parts Nonstructural parts, such as junction boxes, heater ducts, relay shields, and other electrical applications, are manufactured from mat-molded fiber glass. Molded mat fiber glass is short chopped fibers molded in a mat form. The assemblies are fabricated by a process wherein the chopped fibers are molded around a form, bonded together by use of a resin, and cured while under heat and pressure.
Carelessness in removing or handling mat-molded parts can cause the assembly to become damaged.
Vibration may be another factor in the cause of cracks in the assemblies. Damage to mat-molded parts usually consists of holes or cracks (figure 5-101). Similar repair procedures are used for either type of damage.
*-in. dia. max.
2-in. max.
*-in. max.
FIcURE 5-101. Typical damage to mat molded parts.
Repair Procedures The following procedures are typical of those used in the repair of a mat-molded assembly. However, they are not to be construed as the only procedures that could be used. The correct section of the structural repair manual for the specific aircraft should be consulted and followed in all situations.
(1) Inspect the part for location of the crack.
(2) Remove the paint or protective coating from around the damaged ares.
(3) Stop-drill the end of the crack. The size of the twist drill should be not smaller than ⅛ in. and not larger than 18 in.
(4) Lay out and sand the damaged area to the dimensions given in figure 5-102. Remove one-third of the material from both sides of the damaged area. Bevel the area 15° to 45°, as shown in figure 5-102, and sand ½ in. beyond the beveled area.
Beveled edge 3/4 R Edge of sanded area -1/2"
• 3/4" 3/16" stop drill hole - 1/3 T - 1/3 T FIcURE 5-102. Mat molded repair.
(5) Prepare two pieces of PVA (polyviny.
alcohol) film large enough to cover the repair area.
(6) Prepare two pieces of metal large enough to cover the area. Use any piece of metal that will provide satisfactory holding strength.
(7) Check and start the air-circulating furnace. Set the temperature regulator at 220° F.
(8) Select and prepare the resin mixture.
(9) Cut the mat fiber glass material and saturate it in the prepared resin. Cut enough pieces of material to build up the beveled out area to its original contour.
(10) Insert the saturated mat fiber glass material into the repair area. (See figure
5-103.)
Mat fiber glass material impregnated with resin 1/3 T 1/3 T 1/3 T FIcURE 5-103. Insertion of precut saturated sections.
(a) Do one side at a time.
(b) Cover each side with the precut PVA film.
(c) Place the prepared metal plate on each side of the repair as it is completed.
(d) Secure the repair in place by use of C-clamps.
(11) Place the assembly in the preheated oven for at least 1 hr. (check applicable or manufacturer's instructions for the resin used).
(12) Remove the assembly from the oven and let it cool at room temperature.
(13) Disassemble the repair, removing C- clamps, metal plates, and PVA film.
(14) Sand both sides to a smooth finish and to the original contour of the part.
(15) Inspect the repair for soundness, using a metallic ring test. A good repair, when struck with a coin or light aluminum hammer, should resound with a metallic ring.
RADOMES The protective dome or domelike covering for a radar antenna or other radar equipment is called a radome. It must be able to withstand the effects of hail, icing, wind, temperature changes, static electricity, supersonic speeds, and stratospheric altitudes. Also, it must have excellent dielectric qualities.
Handling, Installation, and Storage.
Caution should always be used when handling, installing, or storing aircraft radomes. The necess.
ity for the utmost care to prevent damage to sandwich parts cannot be overemphasized. Radomes (radar and radio antenna housings) are especially susceptible to damage. Damage is sometimes minute and invisible, but when exposed to vibration, stress, or liquids (water or oil), deterioration follows. Microwave distortion and energy losses occur as a result of cracks, punctures, and other physical damage, including moisture and oil contamination.
Take care also to avoid contamination with paint removers and stripping compounds normally used on the metal parts of the aircraft for removing finishes. Some of these materials have been found to penetrate the plastic facings of the radome and may have an adverse effect on its electric properties or its strength. Mild soap and water are used for general cleaning of radome surfaces. When a solvent cleaner is required for removing oils and greases on radome surfaces, use a clean cloth dampened with MEK.
Radomes must be handled with special care. Placing radomes upon rough surfaces and among metal parts must be avoided. Caution should always be used to avoid radome damage resulting from the radome striking against work stands, being dropped, or being drug across rough surfaces.
Correct radome installation begins with the uncrating procedure. Before uncrating a radome, provide a clean padded surface at least as large as that which the radome will occupy when uncrated. Adhere closely to the instructions for opening the radome crate. This will prevent damages that would be inflicted to the radome by protruding nails, bolts, staples, or other sharp objects. Installation instructions outlined in the applicable aircraft maintenance manual must be followed closely when installing radomes. Should sanding or grinding of the radome be required to fit a mounting frame, the sanded surfaces should be classified as a class I repair (discussed later in this chapter) and reworked accordingly.
Radomes should be stored where they are not subject to high humidity. They should be stored in suitable crates or padded racks and supported from the mounting holes. Avoid stacking radomes directly upon each other.
Detection and Removal of Oil and Moisture All radomes are susceptible to moisture and oil contamination. Either can be the cause of very serious degradation of the performance of the aircraft's radar system. Contamination also causes weakening of the radome facing and the facing-to-core bonding strength.
Radomes should be inspected for moisture or oil contaimination prior to repairing or identifying as serviceable; they must be clean and dry prior to electrical testing. Radomes can be checked for mois.
ture pockets using an electronic moisture meter. The probe unit of the meter should be held in contact with the inner surface of the radome and slowly moved over the surface. The presence of moisture will be indicated on the calibrated meter dial. Moisture detection and removal procedures should be accomplished on all radomes before performing repairs.
Inspection For Damage Radomes should be visually inspected for delaminations, scars, scratches, or erosion of the protec- 223
tive coating that would affect only the outer ply.
They should be inspected also for punctures, contamination, fractures of plies affecting either the plies on one side, the plies and core material, or damage extending completely through the outer plies, core material, and inner plies. Different aircraft have different limitations on damage that is reparable, the type of repairs allowed and on damage that is nonreparable. This information can usually be found in the maintenance manuals for the specific aircraft.
Damages to sandwich parts are divided into groups or classes according to the severity and possible effect upon the structure of the aircraft and upon electrical efficiency. Damages are classified in three basic classes: (1) Class I repairs-scars, scratches, or erosion affecting the outer ply only;
(2) class Il repairs punctures, delaminations, contaminations, or fractures in one facing only, possibly accompanied by damage to the core; and (3) class III repairs damage extending completely through the sandwich aflecting both the facings and the core.
Radome Repairs Repair procedures are developed with the objective of equaling as nearly as possible the electrical and strength properties of the original part with a minimum of increase in weight. This can be accom plished by repairing damaged parts with approved materials and working techniques. Therefore radome repairs should be accomplished in accordance with manufacturer's procedures by specially trained personnel of a shop which has proper facilities and adequate test equipment to ensure a satisfactory repair.
Testing of Repairs Radomes must be repaired in a manner that will ensure not only the structural integrity of the radome, but the electrical characteristics as well.
The type of electrical test required after a repair is completed depends on the purpose for which the radome was designed. Typical of the type of electrical tests conducted are:
(1) Transmissivity, which is the average oneway power transmission through the radome or the ratio of power transmitted through the radome to the same power transmitted with the radome removed.
(2) Incidence reflection, the power reflected into the radar system by the radome.
(3) Deflection or refraction to check for possible ghosts or false target returns.
WOODEN AIRCRAFT STRUCTURES While the trend is undoubtedly toward all-metal aircraft, many airplanes still exist in which wood was used as the structural material. The inspection and repair of these wooden structures will continue to be the responsibility of the airframe mechanic.
The ability to inspect wood structures and recognize defects such as dry rot, compression failures, etc., must be developed.
The information in this section is of a general nature and should not be regarded as a substitute for specific instructions contained in the aircraft manufacturer's maintenance and repair manuals.
Methods of construction vary with different types of aircraft, as will the various repair and maintenance procedures.
INSPECTION OF WOODEN STRUCTURES Whenever possible, the aircraft should be kept in a dry, well-ventilated hangar with all inspection covers, access panels, etc., removed for as long as possible prior to inspection. If the aircraft is thoroughly dried out, this will facilitate the inspection, especially when determining the condition of glued Before beginning a detailed inspection of the glued joints and the wood, a rough assessment of the general condition of the structure can sometimes be obtained from the external condition of the aircraft.
The wings, fuselage, and empennage should be checked for undulation, warping or any other departures from the original shape. In instances where the wings, fuselage, or empennage structures and skins form stressed structures (figure 5-104), no departure from the original contour or shape is permissible.
Where light structures using single plywood covering are concerned, some slight sectional undulation or bulging between panels may be permissible provided the wood and glue are sound. However, where such conditions exist, a careful check must be made of the attachment of the ply to its supporting structure. A typical example of a distorted single plywood structure is illustrated in figure 5-105.
The contours and alignment of leading and trailing edges are of particular importance, and a careful check should be made for departure from the original shape. Any distortion of these light ply- 224
Ply skins FIcURE 5-104. Cross sectional view of a stressed-skin structure.
FIGURE 5-105. Single ply structure.
wood and spruce structures is indicative of deterioration, and careful internal inspection will have to be made for security of these parts to the main wing structure, and for general deterioration of the light plywood and spruce members. If deterioration is found in these components, the main wing structure may also be affected.
Splits in the fabric covering on plywood surfaces should not be repaired by doping on another piece of fabric over the affected area. In all cases, the defective fabric should be removed to ascertain whether the plywood skin beneath is serviceable, since it is common for a split in the plywood skin to be responsible for initiating a similar defect in the protective fabric covering.
Although a preliminary inspection of the external structure can be useful in assessing the general condition of the aircraft, it should be noted that wood and glue deterioration can often take place inside a structure without any external indications.
Where moisture can enter a structure, it will seek the lowest point where it will stagnate and promote rapid deterioration. It should also be noted that glue deterioration can take place through other causes without the presence of water.
Glue failure and wood deterioration are often closely allied, and the inspection of glued joints must include an examination of the adjacent wood structure.
The inspection of a complete aircraft for glue or wood deterioration will necessitate checks on parts of the structure which may be known or suspected trouble spots and which are, in many instances, boxed in or otherwise inaccessible. In such instances, considerable dismantling is required, and it may be necessary to cut access holes in ply structures to facilitate the inspection. Such work must be done only in accordance with approved drawings or the repair manual for the aircraft concerned.
Glued Joint Inspection The inspection of glued joints in aircraft structures presents considerable difficulties. Even where access to the joint exists, it is still difficult to positively assess the integrity of the joint, Keep this in mind when inspecting wooden structures.
Some of the more common factors which may cause glue deterioration are: (1) Chemical reactions of the glue caused by aging or moisture, extreme temperatures, or a combination of these factors, (2) mechanical forces caused mainly by wood shrinkage, and (3) development of fungus growths.
Aircraft exposed to large cyclic changes of temperature and humidity are especially prone to wood shrinkage which may lead to glue deterioration.
The amount of movement of wooden members due to these changes varies with the size of each member, the rate of growth of the tree from which the wood was cut and the way in which the wood was converted. Thus, two major members in an aircraft 225
structure, secured to each other by glue, are unlikely to have identical characteristics. Differential loads will, therefore, be transmitted across the glue film since the two members will not react in an identical manner relative to each other. This will impose stresses in the glued joint which can normally be accommodated when the aircraft is new and for some years afterwards. However, glue tends to deteriorate with age, and stresses at the glued joints may cause failure of the joints. This is true even when the aircraft is maintained under ideal conditions.
When checking a glue line (the edge of the glued joint) for condition, all protective coatings of paint should be removed by careful scraping. It is important to ensure that the wood is not damaged during the scraping operation. Scraping should cease immcdiately when the wood is revealed in its natural state and the glue line is clearly discernible.
The glue line is often inspected by the use of a magnifying glass. Where the glue line tends to part or where the presence of glue cannot be detected or is suspect, the glue line should be probed with a thin feeler gage. If any penetration is possible, the joint should be regarded as defective. It is impor tant to ensure that the surrounding wood is dry;
otherwise, a false impression of the glue line would be obtained due to closing of the joint by swelling.
In instances where pressure is exerted on a joint, either by the surrounding structure or by metal Metal fitting Bolt attachment devices such as bolts or screws, a false impression of the glue condition could be obtained unless the joint is relieved of this pressure before the glue line inspection is carried out.
The choice of feeler gage thickness will vary with the type of structure, but a rough guide is that the thinnest possible gage should be used. Figure 5-106 indicates the points where checks with a feeler gage should be made.
Wood Condition Dry rot and wood decay are not usually difficult to detect. Dry rot is indicated by small patches of crumbling wood. A dark discoloration of the wood surface or gray stains running along the grain are indicative of water penetration. If such discoloration cannot be removed by light scraping, the part should be replaced. Local staining of the wood by the dye from a synthetic adhesive hardener can be disregarded.
In some instances where water penetration is sus.
pected, the removal of a few screws from the area in question will reveal, by their degree of corrosion, the condition of the surrounding joint (figure 5-107).
The adhesive will cause slight corrosion of the screw following the original construction; there fore, the condition of the screw should be compared with that of a similar screw removed from another Ply spar web Laminated spar
• Shrinkage Inspection hole in web At all points marked"A", check Shrinkage for glue condition and separation.
FIGURE 5-106. Laminated joint.
Fuselage inner and outer ply skins Position to check Reinforced laminated for separation fuselage member Bulkhead frame Screw hole member Woodscrew Bulkhead ply web Expansion gap (not to be confused with joint separation) Corrosion indicating failure of bulkhead glued joint to fuselage side FicuRE 5-107. Checking a wooden structure for water penetration.
part of the structure known to be free from water soakage.
Plain brass screws are normally used for reirforcing glued wooden members, although zinccoated brass is sometimes used. For hardwoods such as mahogany or ash, steel screws are sometimes used. Unless otherwise specified by the aircraft manufacturer, it is usual to replace screws with new screws of identical length but one gage larger.
Another means of detecting water penetration is to remove the bolts holding the fittings at spar root-end joints, aileron hinge brackets, etc. (figure 5-107). Corrosion on the surface of such bolts and wood discoloration will provide a useful indication of water penetration.
Experience with a particular aircraft will indicate those portions of the structure most prone to water penetration and moisture entrapment, e.g., at windows or the bottom lower structure of entrance doors. However, this is not necessarily indicative of the condition of the complete aircraft.
The condition of the fabric covering on ply surfaces is of great importance. If any doubt exists regarding its protective qualities or if there are any signs of poor adhesion, cracks, or other damage, it should be removed to reveal the ply akin.
The condition of the exposed ply surface should be examined. Water penetration will be shown by dark gray streaks along the grain and a dark discoloration at ply joints or screw holes. If these marks cannot be removed by light scraping or, in the case 227
of advanced deterioration, where there are small surface cracks or separation of the ply laminations, the ply should be replaced. Where evidence of water penetration is found, sufficient areas of the ply surface should be uncovered to determine its extent.
During the inspection, the structure should be examined for other defects of a mechanical nature.
Information regarding such defects is presented in the following paragraphs.
Where bolts secure fittings which take load-carry.
ing members, or where the bolts are subject to landing or shear loads, the bolt holes should be examined for elongation or surface crushing of the wood fibers. The bolts should be removed to facilitate the inspection. It is important to ensure that the bolts are a good fit in the holes.
Check for evidence of damage such as bruises or crushing of structural members which can be caused, for example, by overtightening bolts. Repair techniques for such damage are governed by the extent and depth of the defect.
Compression failures, often wrongly referred to as compression "shakes," are caused by rupture across the wood fibers. This is a serious defect which, at times, is difficult to detect. Special care is necessary when inspecting any wooden member which has been subjected to abnormal bending or compression loads during a hard landing. In the case of a member having been subjected to an excessive bending load, the failure will appear on the surface which has been compressed. The surface subjected to tension will normally show no defects.
In the case of a member taking an excessive direct compression load, the failure will usually be apparent on all surfaces.
If a compression failure is suspected, a flashlight shone along the member, with the beam of light running parallel to the grain, will assist in revealing this type of failure.
A glued joint may fail in service as a result of an accident or because of excessive mechanical loads having been imposed upon it, either in tension or in shear. It is often difficult to decide the nature of the load which caused the failure, but remember that glued joints are generally designed to take shear loads.
If a glued joint is known to have failed in tension, it is difficult to assess the quality of the joint, since these joints may often show an apparent lack of adhesion. Tension failures often appear to strip the glue from one surface leaving the bare wood. In such cases, the glue should be examined with a magnifying glass, which should reveal a fine layer of wood fibers on the glued surface. The presence of fibers indicates that the joint itself is not at fault.
If examination of the glue under magnification does not reveal any wood fibers but shows an imprint of the wood grain, this is caused by pre-drying of the glue before applying pressure during the manufacture of the joints. If the glue exhibits an irregular appearance with star-shaped patterns, this is an indication that pre-curing has occurred before pressure was applied or that pressure has been incorrectly applied or maintained. In all such instances other joints in the aircraft should be suspect.
If a joint is expected to take tension loads, it will be secured by a number of bolts or screws (or both) in the area of tension loading. If a failure occurs in this area, it is usually very difficult to form an opinion of the actual reasons for it, because of considerable breakup of the wood close to the bolts.
In all cases of glued joint failure, whatever the direction of loading, there should be a fine layer of wood fibers adhering to the glue, whether or not the glue has come away completely from one section of the wood member. If there is no evidence of fiber adhesion, this may indicate glue deterioration.
SERVICE AND REPAIR OF WOODEN STRUCTURES Damage to wooden structures such as wing ribs, spars, and skin frequently requires repair. Whenever major wood parts have been damaged, a detailed inspection must be made. Secondary cracks sometimes start some distance away from the main damage and proceed in unrelated directions.
The purpose of repairing all wooden structural parts is to obtain a structure as strong as the original. Severe damage will require replacement of the entire damaged assembly, but minor damage can be repaired by cutting away the damaged members and replacing them with new sections. This replacement is accomplished by glued, or sometimes glued and nailed, or glued and screw-reinforced splicing.
Materials Several forms of wood are commonly used in aircraft. Solid wood or the adjective "solid" used with such nouns as beam or spar refers to a member consisting of one piece of wood.
Laminated wood is an assembly of two or more layers of wood which have been glued together with the grain of all layers or laminations approximately parallel. Plywood is an assembled product of wood and glue that is usually made of an odd number of 228
thin plies (veneers) with the grain of each layer at an angle of 90° with the adjacent ply or plies.
High-density material includes compreg, impreg, or similar commercial products, heat stabilized wood, or any of the hardwood plywoods commonly used as bearing or reinforcement plates. The woods listed in figure 5-108 are those used for structural purposes. For interior trim, any of the decorative woods such as maple or walnut can be used since strength is of little consideration in this situation.
Strength Maximum properties as permissible Species of wood compared to grain devi- Remarks spruce ation (slope of grain) Spruce 100% 1:15 Excellent for all uses. Considered as standard for this table.
Douglas Fir Exceeds spruce 1:15 May be used as substitute for spruce in same sizes or in slightly reduced sizes providing reductions are substantiated. Diticult to work with hand tools. Some tendency to split and splinter during fabrication. Large solid pieces should be avoided due to inspection difficulties.
Gluing satisfactory:
Noble Fir Slightly exceeds spruce 1:15 Satisfactory characteristics with respect to except & percent de- workability, warping, and splitting. May De ficient in shear. used as direct substitute for spruce in same sizes providing shear does not become critical.
Hardness somewhat less than spruce. Gluing satisfactory.
Western Hemlock Slightly exceeds spruce 1:15 Less unitorn in texture than spruce. May t used as direct substitute for spruce. Gluing satisfactory.
Pine, Northern White Properties between 85 1:15 Excellent working qualities and uniform in percent and 96 per- properties but somewhat low in hardness and cent those of spruce. shock-resisting capacity. Cannot be used as substitute for spruce without increase in sizes to compensate for lesser strength. Gluing satisfactory.
White Cedar, Port Orford Exceeds spruce 1:15 May be used as substitute for spruce in same sizes or in slightly reduced sizes providing reductions are substantiated. Easy to work with hand tools. Gluing difficult but satisfactory joints can be obtained if suitable precautions are taken.
Poplar, Yellow Slightly less than spruce 1:15 Excellent working qualities. Should not be except in compression used as a direct substitute for spruce without (crushing) and shear. carefully accounting for slightly reduced strength properties. Somewhat low in shock-resisting capacity. Gluing satisfactory.
FIGURE 5-108. Woods for aircraft use.
All wood and plywood used in the repair of aircraft structures must be of aircraft quality. The species used to repair a part should be the same as that of the original whenever possible. If it is necessary to substitute a different species, follow the recommendations of the aircraft manufacturer in selecting a different species.
Defects Permitted.
a. Cross grain. Spiral grain, diagonal grain, or a combination of the two is acceptable providing the grain does not diverge from the longitudinal axis of the material more than specified in column 3 of figure 108. A check of all four faces of the board is necessary to determine the amount of 229