METAL STRUCTURE, WELDING, AND BRAZING

CHAPTER 4. METAL STRUCTURE, WELDING, AND BRAZING SECTION 1. IDENTIFICATION OF METALS

4-1. GENERAL

Proper identification of the aircraft structural material is the first step in ensuring that the continuing airworthiness of the aircraft will not be degraded by making an improper repair using the wrong materials.

a. Ferrous (iron) alloy materials are generally classified according to carbon content. (See Table 4-1.)

4-2. IDENTIFICATION OF STEEL STOCK

The Society of Automotive Engineers (SAE) and the American Iron and Steel Institute (AISI) use a numerical index system to identify the composition of various steels. The numbers assigned in the combined listing of standard steels issued by these groups represent the type of steel and make it possible to readily identify the principal elements in the material.

a. The basic numbers for the four digit series of the carbon and alloy steel may be found in Table 4-2. The first digit of the four number designation indicates the type to which the steel belongs. Thus, “1” indicates a carbon steel, “2” a nickel steel, “3” a nickel chromium steel, etc. In the case of simple alloy steels, the second digit indicates the approximate percentage of the predominant alloying element. The last two digits usually indicate the mean of the range of carbon content. Thus, the designation “1020” indicates a plain carbon steel lacking a principal alloying element and containing an average of 0.20 percent (0.18 to 0.23) carbon. The designation “2330” indicates a nickel steel of approximately 3 percent (3.25 to 3.75) nickel and an average of 0.30 percent (0.28 to 0.33) carbon content. The designation “4130” indicates a chromium molybdenum steel of approximately 1 percent (0.80 to 1.10) chromium, 0.20 percent (0.15 to 0.25) molybdenum, and 0.30 percent (0.28 to 0.33) carbon.

b. There are numerous steels with higher percentages of alloying elements that do not fit into this numbering system. These include a large group of stainless and heat resisting alloys in which chromium is an essential alloying element. Some of these alloys are identified by three digit AISI numbers and many others by designations assigned by the steel company that produces them. The few examples in Table 4-3 will serve to illustrate the kinds of designations used and the general alloy content of these steels.

c. “1025” welded tubing as per SAE Specification AMST5066B and “1025” seamless tubing are interchangeable.

4-3. INTERCHANGEABILITY OF STEEL TUBING

4-4. IDENTIFICATION OF ALUMINUM

4-5.–4-15. [RESERVED.]

SECTION 2. TESTING OF METALS

4-16. HARDNESS TESTING

If the material type is known, a hardness test is a simple way to verify that the part has been properly heat-treated. Hardness testers such as Rockwell, Brinell, and Vickers can be useful to check metals for loss of strength due to exposure to fire or abusive heating. Also, understrength bolts can be found and removed from the replacement part inventory by checking the hardness of the bolt across the hex flats. Although hardness tests are generally considered nondestructive, hardness testing does leave a small pit in the surface; therefore, hardness tests should not be used on sealing surfaces, fatigue critical parts, load bearing areas, etc., that will be returned to service. These hardness tests provide a convenient means for determining, within reasonable limits, the tensile strength of steel. It has several limitations in that it is not suitable for very soft or very hard steels. Hardness testing of aluminum alloys should be limited to distinguishing between annealed and heat-treated material of the same aluminum alloy. In hardness testing, the thickness and the edge distance of the specimen being tested are two factors that must be considered to avoid distortion of the metal. Several readings should be taken and the results averaged. In general, the higher the tensile strength, the greater its hardness. Common methods of hardness testing are outlined in the following paragraphs. These tests are suitable for determining the tensile properties resulting from the heat treatment of steel. Care should be taken to have case-hardened, corroded, pitted, decarburized, or otherwise nonuniform surfaces removed to a sufficient depth. Exercise caution not to cold-work, and consequently harden, the steel during removal of the surface.

4-17. ROCKWELL HARDNESS TEST

The Rockwell hardness test is the most common method for determining hardness of ferrous and many nonferrous metals. (See Table 4-5.) It differs from Brinell hardness testing in that the hardness is determined by the depth of indentation made by a constant load impressing on an indenter. In this test, a standard minor load is applied to set a hardened steel ball or a diamond cone in the surface of the metal, followed by the application of a standard major load. The hardness is measured by depth of penetration. Rockwell superficial hardness tests are made using light minor and major loads and a more sensitive system for measuring depth of indentation. It is useful for thin sections, very small parts, etc. Calibration of Rockwell hardness testers is done in accordance with ASTM International (ASTM) E-18 specifications.

4-18. BRINELL HARDNESS TEST

In this test, a standard constant load (usually 500 kg to 3,000 kg) is applied to a smooth flat metal surface by a hardened steel-ball type indenter 10 mm in diameter. The 500-kg load is usually used for testing nonferrous metals such as copper and aluminum alloys, whereas the 3,000-kg load is most often used for testing harder metals such as steels and cast irons. The numerical value of Brinell Hardness (HB) is equal to the load divided by the surface area of the resulting spherical impression.

Where:

P=load in kg

D=diameter of the ball in mm d=diameter of the indentation in mm

a. General Precautions. To avoid misapplication of Brinell hardness testing, the fundamentals and limitations of the test procedure must be clearly understood. To avoid inaccuracies, the following rules should be followed:

b. Limitations. The Brinell hardness test has three principal limitations:

c. Calibration. A Brinell hardness tester should be calibrated to meet ASTM standard E10 specifications.

4-19. VICKERS HARDNESS TEST

In this test, a small pyramidal diamond is pressed into the metal being tested. The Vickers Hardness number (HV) is the ratio of the load applied to the surface area of the indention. This is done with the following formula:

4-20. MICROHARDNESS TESTING

This is an indentation hardness test made with loads not exceeding 1 kg (1,000 g). Such hardness tests have been made with a load as light as 1 g, although the majority of microhardness tests are made with loads of 100 g to 500 g. In general, the term is related to the size of the indentation rather than to the load applied.

a. Fields of Application. Microhardness testing is capable of providing information regarding the hardness characteristics of materials which cannot be obtained by hardness tests such as the Brinell or Rockwell, and are as follows:

b. Indenters. Microhardness testing can be performed with either the Knoop or the Vickers indenter. The Knoop indenter is used mostly in the United States; the Vickers indenter is more widely used in Europe.

Where:

P=the applied load in kg

A=the unrecovered projected area in square mm l=the measured length of the long diagonal in mm C=0.07028 (a constant of the indenter, relating projected area of the indentation to the square of the length of the long diagonal)

4-21. INDENTATIONS

The Vickers indenter penetrates about twice as far into the work piece as does the Knoop indenter. The diagonal of the Vickers indentation is about one-third of the total length of the Knoop indentation. The Vickers indenter is less sensitive to minute differences in surface conditions than is the Knoop indenter. However, the Vickers indentation, because of the shorter diagonal, is more sensitive to errors in measuring than is the Knoop indentation. (See Figure 4-1.)

4-22. MAGNETIC TESTING

Magnetic testing consists of determining whether the specimen is attracted by a magnet. Usually, a metal attracted by a magnet is iron, steel, or an iron-base alloy containing nickel, cobalt, or chromium. However, there are exceptions to this rule since some nickel and cobalt alloys may be either magnetic or nonmagnetic. Never use this test as a final basis for identification. The strongly attracted metals could be pure iron, pure nickel, cobalt, or iron-nickel-cobalt alloys. The lightly attracted metals could be cold-worked stainless steel or Monel. The nonmagnetic metals could be aluminum, magnesium, silver, or copper-base alloy, or an annealed 300-type stainless steel.

4-23. ALUMINUM TESTING

Hardness tests are useful for testing aluminum alloy chiefly as a means of distinguishing between annealed, cold-worked, heat-treated, and heat-treated and aged material. It is of little value in indicating the strength or quality of heat-treatment. Typical hardness values for aluminum alloys are shown in Table 4-5.

4-24.–4-35. [RESERVED.]

SECTION 3. PRECAUTIONARY MEASURES

4-36. FLUTTER AND VIBRATION PRECAUTIONS

To prevent the occurrence of severe vibration or flutter of flight control surfaces during flight, precautions must be taken to stay within the design balance limitations when performing maintenance or repair.

a. Balance Changes. The importance of retaining the proper balance and rigidity of aircraft control surfaces cannot be overemphasized. The effect of repair or weight change on the balance and center of gravity is proportionately greater on lighter surfaces than on the older heavier designs. As a general rule, repair the control surface in such a manner that the weight distribution is not affected in any way, in order to preclude the occurrence of flutter of the control surface in flight. Under certain conditions, counter-balance weight is added forward of the hinge line to maintain balance. Add or remove balance weights only when necessary in accordance with the manufacturer’s instructions. Flight testing must be accomplished to ensure flutter is not a problem. Failure to check and retain control surface balance within the original or maximum allowable value could result in a serious flight hazard.

b. Painting and Refinishing. Special emphasis is directed to the effect of too many extra coats of paint on balanced control surfaces. Mechanics must avoid adding additional coats of paint in excess of what the manufacturer originally applied. If available, consult the aircraft manufacturer’s instructions relative to finishing and balance of control surfaces.

c. Trapped Water or Ice. Instances of flutter have occurred from unbalanced conditions caused by the collection of water or ice within the surface. Therefore, ventilation and drainage provisions must be checked and retained when maintenance is being done.

d. Trim Tab Maintenance. Loose or vibrating trim tabs will increase wear of actuating mechanisms and hinge points which may develop into serious flutter conditions. When this happens, primary control surfaces are highly susceptible to wear, deformation, and fatigue failures because of the buffeting nature of the airflow over the tab mechanism. Trailing-edge play of the tab may increase, creating an unsafe flutter condition. Careful inspection of the tab and its mechanism should be conducted during overhaul and annual inspection periods. Compared to other flight control systems on the aircraft, only a minor amount of tab-mechanism wear can be tolerated.

NOTE: If the pilot has experienced flutter, or thinks they have, then a complete inspection of the aircraft flight control system and all related components, including rod ends, bearings, hinges, and bellcranks, must be accomplished. Suspected parts should be replaced.

4-37. LOAD FACTORS FOR REPAIRS

In order to design an effective repair to a sheet metal aircraft, the stresses that act on the structure must be understood.

4-38. TRANSFER OF STRESSES WITHIN A STRUCTURE

An aircraft structure must be designed in such a way that it will accept all of the stresses imposed upon it by the flight and ground loads without any permanent deformation. Any repair made must accept the stresses, carry them across the repair, and then transfer them back into the original structure. These stresses are considered as flowing through the structure, so there must be a continuous path for them, with no abrupt changes in cross-sectional areas along the way. Abrupt changes in cross-sectional areas of aircraft structure that are subject to cycle loading/stresses will result in stress concentration that may induce fatigue cracking and eventual failure. A scratch or gouge in the surface of a highly stressed piece of metal will cause a stress concentration at the point of damage.

a. Multirow Fastener Load Transfer. When multiple rows of rivets are used to secure a lap joint, the transfer of stresses is not equal in each row. The transfer of stress at each row of rivets may be thought of as transferring the maximum amount capable of being transferred without experiencing rivet shear failure.

b. Use Of Stacked Doublers. A stacked doubler is composed of two or more sheets of material that are used in lieu of a single, thicker sheet of material. Because the stress transferred at each row of rivets is dependent upon the maximum stress that can be transferred by the rivets in that row, the thickness of the sheet material at that row need only be thick enough to transfer the stress applied. Employing this principle can reduce the weight of a repair joint.

4-39.–4-49. [RESERVED.]

SECTION 4. METAL REPAIR PROCEDURES

4-50. GENERAL

The airframe of a fixed-wing aircraft is generally considered to consist of five principal units: the fuselage, wings, stabilizers, flight control surfaces, and landing gear.

4-51. RIVETED (OR BOLTED) STEEL TRUSS-TYPE STRUCTURES

Repairs to riveted structures may be made employing the general principles outlined in the following paragraphs on aluminum alloy structures. Repair methods may also be found in textbooks on metal structures. Methods for repair of the major structural members must be specifically approved by the Federal Aviation Administration (FAA).

4-52. ALUMINUM ALLOY STRUCTURES

Extensive repairs to damaged stressed skin on monocoque-types of aluminum alloy structures must be made in accordance with FAA-approved manufacturer’s instructions or another FAA-approved source.

a. Rivet Holes. Rivet holes are slightly larger than the diameter of the rivet. When driven, solid rivets expand to fill the hole. The strength of a riveted joint is based upon the expanded diameter of the rivet. Therefore, it is important that the proper drill size be used for each rivet diameter.

b. Disassembly Prior to Repairing. If the parts to be removed are essential to the rigidity of the complete structure, support the structure prior to disassembly in such a manner as to prevent distortion and permanent damage to the remainder of the structure. When rivets are removed, undercut rivet heads by drilling. Use a drill of the same size as the diameter of the rivet. Drilling must be exactly centered and to the base of the head only. After drilling, break off the head with a pin punch and carefully drive out the shank. On thin or unsupported metal skin, support the sheet metal on the inside with a bucking bar. Removal of rivet heads with a cold chisel and hammer is not recommended because skin damage and distorted rivet holes will probably result. Inspect rivet joints adjacent to damaged structure for partial failure by removing one or more rivets to see if holes are elongated or the rivets have started to shear.

c. Effective Tools. Care must also be taken whenever screws must be removed to avoid damage to adjoining structure. When properly used, impact wrenches can be effective tools for removal of screws; however, damage to adjoining structure may result from excessive vertical loads applied through the screw axis. Excessive loads are usually related to improperly adjusted impact tools or attempting to remove screws that have seized from corrosion. Remove seized screws by drilling and use of a screw extractor. Once the screw has been removed, check for structural cracks that may appear in the adjoining skin doubler, or in the nut or anchor plate.

4-53. SELECTION OF ALUMINUM FOR REPLACEMENT PARTS

All aluminum replacement sheet metal must be identical to the original or properly altered skin. If another alloy is being considered, refer to the information on the comparative strength properties of aluminum alloys contained in MIL-HDBK-5 or MMPDS-10.

a. Temper. The choice of temper depends upon the severity of the subsequent forming operations. Parts having single curvature and straight bend lines with a large bend radius may be advantageously formed from heat-treated material; while a part, such as a fuselage frame, would have to be formed from a soft, annealed sheet, and heat-treated after forming. Make sure sheet metal parts which are to be left unpainted are made of clad (aluminum coated) material. Make sure all sheet material and finished parts are free from cracks, scratches, kinks, tool marks, corrosion pits, and other defects which may be factors in subsequent failure.

b. Use of Annealed Alloys for Structural Parts. The use of annealed aluminum alloys for structural repair of an aircraft is not recommended. An equivalent strength repair using annealed aluminum will weigh more than a repair using heat-treated aluminum alloy.

4-54. HEAT TREATMENT OF ALUMINUM ALLOY PARTS

All structural aluminum alloy parts are to be heat-treated in accordance with the heat-treatment instruction issued by the manufacturers of the part. In the case of a specified temper, the sequence of heat-treating operations set forth in MIL-HDBK-5 or MMPDS-10 and corresponding specifications. If the heat-treatment produces warping, straighten the parts immediately after quenching. Heat-treat riveted parts before riveting, to preclude warping and corrosion.

a. Quenching. Quench material from the solution heat-treating temperature as rapidly as possible after removal from the furnace. Quenching in cold water is preferred, although less drastic chilling (hot or boiling water, or airblast) is sometimes employed for bulk sections, such as forgings, to minimize quenching stresses.

b. Reheating at Temperatures Above Boiling Water. Reheating of 2017 and 2024 alloys above 212 °F tend to impair the original heat treatment. Therefore, reheating above 212 °F, including the baking of primers, is not acceptable without subsequent complete and correct heat treatment.

4-55. BENDING METAL

When describing a bend in aviation, the term “bend radii” (BR) is used to refer to the inside radius. Requirements for bending the metal to various shapes are frequently encountered. When a metal is bent, it is subjected to changes in its grain structure, causing an increase in its hardness.

Table 4-6. For example, the minimum BR for 0.016 thick 2024-T6 (alloy and temper) is found is

found to be 2 to 4 times the material thickness, or 0.032 to 0.064.

4-56. SETBACK

b. To determine setback for a bend of more or less than 90 degrees, a correction known as a K-factor must be applied to find the setback.

(1) Table 4-7 shows a chart of K-factors. To find the setback for any degree of bend, multiply the sum of the BR and metal thickness by the K-value for the angle through which the metal is bent.

(2) Figure 4-3 shows an example of a piece of 0.064-inch sheet metal bent through 45 degrees to form an open angle of 135 degrees. For 45 degrees, the K-factor is 0.41421. The setback, or the distance from the mold point to the bend tangent line, is:

(3) If a closed angle of 45 degrees is formed, the metal must be bent through 135 degrees. The K-factor for 135 degrees is 2.4142, so the setback, or distance from the mold point to the bend tangent line, is 0.758 inch.

4-57. RIVETING

(3) An example of identification marking of a rivet follows:

• MS 20470AD3-5 Complete part number
• MS Military standard number
• 20470 Universal head rivet
• AD 2117-T aluminum alloy
• 3 3/32 inch in diameter
• 5 5/16 inch in length

d. The 2117 rivets may be driven in the condition received, but 2017 rivets above 3/16 inch in diameter and all 2024 rivets are to be kept packed in dry ice or refrigerated in the “quenched” condition until driven, or be reheat-treated just prior to driving, as they would otherwise be too hard for satisfactory riveting. Dimensions for formed rivet heads are shown in Figure 4-6, item (a), together with commonly found rivet imperfections.

NOTE: Blind rivets shall not be used:

CAUTION: For metal repairs to the airframe, the use of blind rivets must be specifically authorized by the airframe manufacturer or approved by a representative of the FAA.

(1) Self-plugging friction-lock cherry rivets. This patented rivet may be installed when there is access to only one side of the structure. The blind head is formed by pulling the tapered stem into the hollow shank. This swells the shank and clamps the skins tightly together. When the shank is fully upset, the stem pulls in two. The stem does not fracture flush with the rivet head and must be trimmed and filed flush for the installation to be complete. Because of the friction-locking stem, these rivets are very sensitive to vibrations. Inspection is visual, with a loose rivet standing out in the standard “smoking rivet” pattern. Removal consists of punching out the friction-locked stem and then treating it like any other rivet. (See Figure 4-7.)

(3) Bulbed CherryLOCK Rivets. One of the earlier types of mechanical-lock rivets developed were Bulbed CherryLOCK blind rivets. These blind rivets have as their main advantage the ability to replace a solid shank rivet size for size. (See Figure 4-8.)

The rough end of the retained stem in the center on the manufactured head must never be filed smooth, because it will weaken the strength of the lockring and the center stem could fall out. (See Figure 4-8.)

(4) The CherryMAX rivet (see Figure 4-9) uses one tool to install three standard rivet diameters and their oversize counterparts. This makes the use of CherryMAX rivets very popular with many small general aviation repair shops. CherryMAX rivets are available in four nominal diameters: 1/8, 5/32, 3/16, and 1/4 inch; and three oversized diameters. CherryMAX rivets are manufactured with two head styles: universal and countersunk. The CherryMAX rivets consists of five parts: bulbed blind header, hollow rivet shell, locking (foil) collar, driving anvil, and pulling stem. The blind bulbed header takes up the extended shank and forms the bucktail on a CherryMAX rivet stem. Rivet sleeves are made from 5056 aluminum, Monel, and INCONEL 600. The stems are made from alloy steel, corrosion resistant steel (CRES), and INCONEL X-750 stem. CherryMAX rivets have an ultimate shear strength ranging from 50 thousands of pounds per square inch (KSI) to 75 KSI.

(5) An Olympic-Lok rivet (see Figure 4-10) is a light three-piece mechanically locked, spindle-type blind rivet. It carries its stem lock integral to the manufactured head. While installing, the lockring is pressed into a groove on the pulling stem just as the rivet completes drawing the metal together. After installation is completed, never file the stem of an Olympic-Lok rivet, because it will weaken the lockring attachment. The Olympic-Lok fastener is available in three head styles: universal protruding, 100-degree flush countersink, and 100-degree flush shear; and three diameters: 1/8, 5/32, and 3/16 inch. The three diameters are available in eight different alloy combinations of 2017-T4, A-286, 5056, and Monel. Olympic-Lok lock spindles are made from the same material as the sleeves.

(6) Huck rivets (see Figure 4-11) are available in two head styles: protruding and flush. They are available in four diameters: 1/8, 5/32, 3/16, and 1/4 inch. Their diameters are measured in increments of 1/32 inch and lengths are measured in 1/16 inch increments. They are manufactured in three different combinations of alloys: 5056 aluminum sleeve with 2024 aluminum alloy pin, A-286 corrosion-resistant steel sleeve with an A-286 pin, and a Monel 400 sleeve with an A-286 pin. The Huck fastener has the ability to tightly draw up two or more sheets of metal together while being installed. After the take up of the Huck fastener is completed, the lockring is squeezed into a groove on the pulling stem. The anvil or footer (of the installation tool) packs the ring into the groove of the pulling stem by bearing against the lockring.

(7) Common pull-type POP rivets, produced for nonaircraft related applications, are not approved for use on certificated aircraft structures or components.

4-58. REPAIR METHODS AND PRECAUTIONS FOR ALUMINUM STRUCTURE

Carefully examine all adjacent rivets outside of the repair area to ascertain that they have not been harmed by operations in adjacent areas. Drill rivet holes round, straight, and free from cracks. Deburr the hole with an oversize drill or deburring tool. The rivet set used in driving the rivets must be cupped slightly flatter than the rivet head. (See Figure 4-6.) Rivets are to be driven straight and tight, but not overdriven or driven while too hard, since the finished rivet must be free from cracks. Information on special methods of riveting, such as flush riveting, usually may be obtained from manufacturer’s service manuals.

a. Splicing of Tubes. Round or streamline aluminum alloy tubular members may be repaired by splicing. (See Figure 4-12.) Splices in struts that overlap fittings are not acceptable. When solid rivets go completely through hollow tubes, their diameter must be at least one-eighth of the outside diameter of the outer tube. Rivets which are loaded in shear should be hammered only enough to form a small head and no attempt made to form the standard roundhead. The amount of hammering required to form the standard roundhead often causes the rivet to buckle inside the tube. (Correct and incorrect examples of this type of rivet application are incorporated in Figure 4-12.)

b. Repairs to Aluminum Alloy Members. Make repairs to aluminum alloy members with the same material or with suitable material of higher strength. The 7075 alloy has greater tensile strength than other commonly used aluminum alloys such as 2014 and 2024, but is subject to somewhat greater notch sensitivity. In order to take advantage of its strength characteristics, pay particular attention to design of parts to avoid notches, small radii, and large or rapid changes in cross-sectional areas. In fabrication, exercise caution to avoid processing and handling defects, such as machine marks, nicks, dents, burrs, scratches, and forming cracks. Cold straightening or forming of 7075-T6 can cause cracking; therefore, it may be advisable to limit this processing to minor cold straightening.

c. Wing and Tail Surface Ribs. Damaged aluminum alloy ribs either of the stamped sheet-metal type or the built-up type employing special sections, square or round tubing, may be repaired by the addition of suitable reinforcement. (Acceptable methods of repair are shown in Figure 4-13 and Figure 4-14.) These examples deal with types of ribs commonly found in small- and medium-size aircraft. Repair schemes developed by the aircraft manufacturer are acceptable, but any other methods of reinforcement are major repairs and require approved data.

d. Trailing and Leading Edges and Tip Strips. Repairs to wing, control surface trailing edges, leading edges, and tip strips should be made by properly executed and reinforced splices. Acceptable methods of trailing edge repairs are shown in Figure 4-15.

e. Repair of Damaged Skin. In cases where metal skin is damaged extensively, repair by replacing an entire sheet panel from one structural member to the next. The repair seams are to lie along stiffening members, bulkheads, etc.; and each seam must be made exactly the same in regard to rivet size, splicing, and rivet pattern as the manufactured seams at the edges of the original sheet. If the two manufactured seams are different, the stronger one will be copied. (See Figure 4-16 for typical acceptable methods of repairs.)

f. Patching of Small Holes. Small holes in skin panels which do not involve damage to the stiffening members may be patched by covering the hole with a patch plate in the manner shown in Figure 4-16. Flush patches also may be installed in stressed-skin type construction. An acceptable and easy flush patch may be made by trimming out the damaged area and then installing a conventional patch on the underneath side or back of the sheet being repaired. A plug patch plate of the same size and skin thickness as the opening may then be inserted and riveted to the patch plate. Other types of flush patches similar to those used for patching plywood may be used. The rivet pattern used, however, must follow standard practice to maintain satisfactory strength in the sheet.

g. Splicing of Sheets. The method of copying the seams at the edges of a sheet may not always be satisfactory. For example, when the sheet has cutouts, or doubler plates at an edge seam, or when other members transmit loads into the sheet, the splice must be designed as illustrated in the following examples.

3 × 0.032 = 0.096 inch. Use 1/8-inch 2117-T4 (AD) rivets (5/32-inch 2117-T4 (AD) would be satisfactory).

(c) Lay out rivet pattern with spacing not less than shown in Figure 4-5. Referring to Figure 4-5, item (A), it seems that a double row pattern with the minimum spacing will give a total of 40 rivets. However, as only 37 rivets are required, two rows of 19 rivets each equally spaced over the 10 inches will result in a satisfactory splice.

h. Straightening of Stringers or Intermediate Frames. Members which are slightly bent may be straightened cold and examined with a magnifying glass for cracks or tears to the material. Reinforce the straightened part to its original shape, depending upon the condition of the material and the magnitude of any remaining kinks or buckles. If any strain cracks are apparent, make complete reinforcement in sound metal beyond the damaged portion.

i. Local Heating. Do not apply local heating to facilitate bending, swaging, flattening, or expanding operations of heat-treated aluminum alloy members, as it is difficult to control the temperatures closely enough to prevent possible damage to the metal, and it may impair its corrosion resistance.

j. Splicing of Stringers and Flanges. It is recommended that all splices be made in accordance with the manufacturer’s recommendations. If the manufacturer’s recommendations are not available, the typical splices for various shapes of sections are shown in Figures 4-17 through 4-19. Design splices to carry both tension and compression, and use the splice shown in Figure 4-18 as an example illustrating the following principles.

k. Size of Splicing Members. When the same material is used for the splicing members as for the original member, the cross-section area (i.e., the shaded areas in Figure 4-17), of the splice material will be greater than the area of the section element which it splices. The area of a section element (e.g., each leg of an angle or channel) is equal to the width multiplied by the thickness. For example, the bar “B” in Figure 4-18 is assumed to splice the upper leg of the stringer, and the angle “A” is assumed to splice the bulbed leg of the stringer. Since the splice bar “B” is not as wide as the adjacent leg, its thickness must be increased such that the area of bar “B” is at least equal to the area of the upper leg of the stringer.

l. The Diameter of Rivets in Stringers. The diameter of rivets in stringers might preferably be between two and three times the thickness “t” of the leg, but must not be more than 1/4 the width “W” of the leg. Thus, 1/8-inch rivets are chosen in the example in Figure 4-18. If the splices were in the lower surface of a wing, the end rivets would be made the same size as the skin-attaching rivets, or 3/32 inch.

m. The Number of Rivets. The number of rivets required on each side of the cut in a stringer or flange may be determined from standard textbooks on aircraft structures, or may be found in Tables 4-9 through 4-11.

n. Splicing of Intermediate Frames. The same principles used for stringer splicing may be applied to intermediate frames when the following point is considered. Conventional frames of channel or Z sections are relatively deep and thin compared to stringers, and usually fail by twisting or by buckling of the free flange. Reinforce the splice joint against this type of failure by using a splice plate heavier than the frame and by splicing the free flange of the frame with a flange of the splice plate. (See Figure 4-20.) Since a frame is likely to be subjected to bending loads, make the length of splice plate “L” more than twice the width “W2,” and the rivets spread out to cover the plate.

4-59. REPAIRING CRACKED MEMBERS

Acceptable methods of repairing various types of cracks in structural elements are shown in Figures 4-21 through 4-24. The following general procedures apply in repairing such defects.

4-60. STEEL AND ALUMINUM FITTINGS

a. Steel Fittings. Inspect for the following defects.

4-61. CASTINGS

Damaged castings are to be replaced and not repaired unless the method of repair is specifically approved by the aircraft manufacturer or substantiating data for the repair has been reviewed by the FAA for approval.

4-62. SELECTIVE PLATING IN AIRCRAFT MAINTENANCE

Selective plating is a method of depositing metal from an electrolyte to the selected area. The electrolyte is held in an absorbent material attached to an inert anode. Plating contact is made by brushing or swabbing the part (cathode) with the electrolyte-bearing anode.

a. Selective Plating Uses. This process can be utilized for any of the following reasons.

b. Specifications. Selective plating (electrodepositions), when properly applied, will meet the following specifications and standards.

c. General Requirements.

d. Equipment. The power source should operate on either 110 V or 220 V alternating current (AC), 60 Hertz (Hz), single-phase input. It should have a capability to produce direct current (DC) having smooth characteristics with controlled ripple and be able to output a current of at least 25 amp at 0 V to 25 V. Minimum instrumentation of the power source should include a voltmeter, ammeter, and amp-hour meter.

e. Materials. The anodes should be of high-purity dense graphite or platinum-iridium alloys. Do not mix solutions from different suppliers. This could result in contamination.

f. Detail Requirements. On large parts, no area greater than approximately 10 percent of the total area of the part should be plated by this selective plating process. Small parts may be partially or completely plated. Special cases exceeding these limitations should be coordinated with the manufacturer of the plating equipment being used and their recommendations should be followed.

g. Anode Selection. As a general guide, the contact area of the anode should be approximately one-third the size of the area to be plated. When selecting the anode, the configuration of the part will dictate the shape of the anode.

h. Required Amp-Hour Calculation. The selected plating solution has a factor which is equal to the amp-hours required to deposit 0.0001-inch on 1 square inch of surface. Determine the thickness of plating desired on a certain area, and multiply the solution factor times the plating thickness times the area in square inches to determine the amp-hours required. This factor may vary because of temperature, current density, etc.

i. Cleaning. Remove corrosion, scale, oxide, and unacceptable plating prior to processing. Use a suitable solvent or cleaner to remove grease or oil.

k. Plating on Copper and Copper Base Alloys.

n. Plating on Cast Iron and High-Carbon Steels (Steels Heat Treated to 180,000 psi).

NOTE: Where the solution vendor provides substantiating data that hydrogen embrittlement will not result from plating with a particular solution, then a postbake is not required. This substantiating data can be in the form of aircraft industry manufacturer’s process specifications, military specifications, or other suitable data.

NOTE: Acid etching should be avoided, if possible. Where etching is absolutely necessary, it should always be done with reverse current. Use alkaline solutions for initial deposits.

p. Dissimilar Metals and Changing Base. As a general rule, when plating two dissimilar metals, follow the plating procedure for the one with the most steps or activation. If activating steps have to be mixed, use reverse current activation steps prior to DC activation steps.

r. Qualification Tests. All brush-plated surfaces should be tested for adhesion of the electrodeposit. Apply a 1-inch wide strip of 3M Scotch Flatback Masking Tape 250, or an approved equal, with the adhesive side to the freshly plated surface. Apply the tape with heavy hand pressure and remove it with one quick motion perpendicular to the plated surface. Any plating adhering to the tape should be cause for rejection.

s. Personnel Training for Quality Control. Manufacturers of selective-plating equipment provide training in application techniques at their facilities. Personnel performing selective plating must have adequate knowledge of the methods, techniques, and practices involved. These personnel should be certified as qualified operators by the manufacturers of the products used.

4-63.–4-73. [RESERVED.]

SECTION 5. WELDING AND BRAZING

4-74. GENERAL

This section covers weld repairs to aircraft and component parts only. Observe the following procedures when using such equipment as gas tungsten arc welding (GTAW), Gas Metal Arc Welding (GMAW), plasma arc welding, and oxyacetylene gas welding. When repairs of any of these flight-critical parts are required, it is extremely important to make the weld repairs equal to the original weld. Identifying the kind of metal to be welded, identifying the kind of welding process used in building the part originally, and determining the best way to make welded repairs are of utmost importance.

f. Aircraft Welding Qualifications. Four groups of metals a person can be certified and qualified to use are:

4-75. EQUIPMENT SELECTION

Use the welding equipment manufacturer’s information to determine if the equipment will satisfy the requirements for the type of welding operation being undertaken. Disregarding such detailed operating instructions may cause substandard welds. For example, when using GTAW equipment, a weld can be contaminated with tungsten if the proper size electrode is not used when welding with DC reverse polarity. Another example, the depletion of the inert gas supply below the critical level causes a reduction in the gas flow and will increase the danger of atmospheric contamination.

a. Electric welding equipment versatility requires careful selection of the type of current and polarity to be used. Since the composition and thickness of metals are deciding factors, the selection may vary with each specific application. Metals having refractory surface oxide films (i.e., magnesium alloys and aluminum and its alloys), are generally welded with AC, while DC is used for carbon, low alloy, noncorrodible, and heat-resisting steels. General recommendations covering current and polarity are shown in Table 4-12.

b. Oxyacetylene gas equipment is suitable for welding most metals. It is not, however, the best method to use on such materials as stainless steel, magnesium, and aluminum alloys; because of base metal oxidization, distortion, and loss of ductility.

NOTE: If oxyacetylene is used for welding stainless steel or aluminum, all flux must be removed, as it may cause corrosion.

4-76. ACCURATELY IDENTIFY THE TYPE OF MATERIAL TO BE REPAIRED

If positive identification of the material is not possible, contact the aircraft manufacturer or subject the item to a metallurgical laboratory analysis. Before any welding is attempted, carefully consider the weldability of the alloy, since all alloys are not readily weldable. The following steels are readily weldable: plain carbon (of the 1000 series), nickel steel (of the SAE 2300 series), chrome-nickel alloys (of the SAE 3100 series), chrome-molybdenum steels (of the SAE 4100 series), and low nickel-chrome-molybdenum steel (of the SAE 8600 series).

4-77. PREPARATION FOR WELDING

4-78. INSPECTION OF A COMPLETED WELD

Visually inspect the completed weld for the following:

4-79. MICROFISSURES

Cracks in parts and materials can vary from tiny microfissures that are visible only with magnification, to those easily identified by unaided eyes. Microfissures are the worst type of defect for two reasons: they are often hard to detect, and they produce the worst form of notch effect/stress concentration. Once they form, they propagate with repeated applications of stress and lead to early failures. Every possible means should be used to detect the presence of cracks, and ensure their complete removal before welding operations proceed. (See Figure 4-26.)

4-80. NONDESTRUCTIVE TESTING

Nondestructive testing or evaluation is advisable in critical applications. Nondestructive testing methods such as; magnetic particle, liquid penetrant, radiography, ultrasonic, eddy current, and acoustic emission can be used; however, they require trained and qualified people to apply them.

4-81. PRACTICES TO GUARD AGAINST

Do not file or grind welds in an effort to create a smooth appearance, as such treatment causes a loss of strength. Do not fill welds with solder, brazing metal, or any other filler. When it is necessary to weld a joint which was previously welded, remove all of the old weld material before rewelding. Avoid welding over a weld, because reheating may cause the material to lose its strength and become brittle. Never weld a joint which has been previously brazed.

4-82. TORCH SIZE (Oxyacetylene welding)

When using oxyacetylene welding, the torch tip size depends upon the thickness of the material to be welded. Commonly used sizes, proven satisfactory by experience, are shown in Table 4-13.

4-83. WELDING RODS AND ELECTRODES

Use welding rods and electrodes that are compatible with the materials to be welded. Welding rods and electrodes for various applications have special properties suitable for the application intended. Lap welds are used in shear applications. The weld throat of the fillet weld is considered the plane 45 degrees to the surface plane of the sheet being welded and is equal to 0.707 times the thickness of the sheet stock. (See Figure 4-27.)

4-84. ROSETTE WELDS

Rosette welds are generally employed to fuse an inner reinforcing tube (liner) with the outer member. Where a rosette weld is used, drill a hole, (in the outside tube only) of sufficient size to insure fusion of the inner tube. A hole diameter of approximately onefourth the tube diameter of the outer tube serves adequately for this purpose. In cases of tightfitting sleeves or inner liners, the rosettes may be omitted. Rosette weld edge distance is 1/2 the diameter of the tube, as measured from the edge of the rosette hole to the end of the inside and outside tube. Rosettes shall not be considered when determining the strength of a welded form. Drill an 1/8-inch hole in the lower tube in the center of the intended rosette weld so the heat does not burn away the outer tube. This small hole tends to bleed off the heat from the torch and keeps the size of the rosette small.

4-85. HEAT-TREATED MEMBERS

Certain structural parts may be heat treated and, therefore, could require special handling. In general, the more responsive an alloy steel is to heat treatment, the less suitable it is for welding because of its tendency to become brittle and lose its ductility in the welded area. Weld the members which depend on heat treatment for their original physical properties by using a welding rod suitable for producing heat-treated values comparable to those of the original members. (See paragraph 4-74.) After welding, heat treat the affected members to the manufacturer’s specifications.

4-86. TYPES OF WELDING

a. Gas Welding. A fuel gas such as acetylene or hydrogen is mixed inside a welding torch with oxygen to produce a flame with a temperature of around 6,300 °F (3,482 °C). This flame is used to melt the materials to be welded. A filler rod is melted into the puddle of molten metal to reinforce the weld. When highly reactive metals such as aluminum are gas welded, they must be covered with flux to exclude oxygen from the molten metal and keep oxides from forming which would decrease the strength of the weld. (An illustration of a carburizing flame, a neutral flame, and an oxidizing flame is shown in Figure 4-28.)

b. Shielded Metal Arc Welding (SMAW). This method is the most familiar and common type and is known in the trade as stick welding. A metal wire rod coated with a welding flux is clamped in an electrode holder connected to the power supply with a heavy electrical cable. The metal to be welded is also attached to the power supply. The electrical power is supplied to the work at a low voltage and high current and may be either AC or DC, depending upon the type of welding being done. An arc is struck between the rod and the work and produces heat in excess of 10,000 °F, which melts both the material and the rod. As the flux melts, it releases an inert gas which shields the molten puddle from oxygen in the air and prevents oxidation. The molten flux covers the weld and hardens to an airtight slag cover that protects the weld bead as it cools. This slag must be chipped off to examine the weld.

c. Gas Metal Arc Welding (GMAW). This method of welding was formerly called Metal Inert Gas (MIG) welding, and is an improvement over stick welding because an uncoated wire electrode is fed into the torch and an inert gas such as argon, helium, or carbon dioxide flows out around the wire to protect the puddle from oxygen. The power supply connects between the torch and the work, and the arc produces the intense heat needed to melt the work and the electrode. Low-voltage high-current DC is used almost exclusively with GMAW welding. GMAW is used more for large-volume production work than for aircraft repair.

d. Gas Tungsten Arc Welding (GTAW). This is the form of electric arc welding that fills most of the needs in aircraft maintenance. It is more commonly known as tungsten inert gas (TIG) welding.

4-87. ELECTRIC-RESISTANCE WELDING

Many thin sheet metal parts for aircraft, especially stainless steel parts, are joined by one of the forms of electric resistance welding, either spot welding or seam welding.

a. Spot Welding. Two copper electrodes are held in the jaws of the spot welding machine, and the material to be welded is clamped between them. Pressure is applied to hold the electrodes tightly together, and electrical current flows through the electrodes and the material. The resistance of the material being welded is so much higher than that of the copper electrodes that enough heat is generated to melt the metal. The pressure on the electrodes forces the molten spots in the two pieces of metal to unite, and this pressure is held after the current stops flowing long enough for the metal to solidify. Refer to MIL-HDBK-5 or MMPDS-10 for joint construction and strength data. The amount of current, pressure, and dwell time are all carefully controlled and matched to the type of material and the thickness to produce the correct spot welds. (See Figure 4-31.)

b. Seam Welding. Rather than having to release the electrodes and move the material to form a series of overlapping spot welds, a seam-welding machine is used to manufacture fuel tanks and other components where a continuous weld is needed. Two copper wheels replace the bar-shaped electrodes. The metal to be welded is moved between them, and electric pulses create spots of molten metal that overlap to form the continuous seam.

4-88. BRAZING

Brazing refers to a group of metal-joining processes in which the bonding material is a nonferrous metal or alloy with a melting point higher than 425 ºC (800 °F), but lower than that of the metals being joined. Brazing includes silver brazing (erroneously called silver soldering or hard soldering), copper brazing, and aluminum brazing.

NOTE: Never weld over a previously brazed joint.

(2) Alloys commonly referred to as silver solders melt above 425 °C (800 °F), and when using them the process should be called silver brazing.

4-89. AIRCRAFT PARTS NOT TO BE WELDED

a. Brace Wires and Cables. Do not weld aircraft parts whose proper function depends upon strength properties developed by cold-working. Among parts in this classification are streamlined wire and cables.

b. Brazed and Soldered Parts. Do not weld brazed or soldered parts as the brazing mixture or solder will penetrate and weaken the hot steel.

c. Alloy Steel Parts. Do not weld alloy steel parts such as aircraft bolts, turnbuckle ends, etc., which have been heat treated to improve their mechanical properties.

d. Nos. 2024 and 7075 Aluminum. Do not weld these two aluminum alloys (that are often used in aircraft construction) because the heat from the welding process will cause severe cracking. The 2024 aluminum is most often used in wing skins, fuselage skins, and in most structured airframe parts. The 7075 aluminum is most often used in machined fittings such as wing spar attachments, landing-gear attachments, and other structural parts.

4-90. WELDING ROD SELECTION

Most aircraft repair shops that are prepared to make weld repairs should have the basic selection of welding rods available. The best rods to stock, the metals they weld, and the AWS specification numbers are shown in Table 4-15.

4-91. REPAIR OF TUBULAR MEMBERS

a. Inspection. Prior to repairing tubular members, carefully examine the structure surrounding any visible damage to insure that no secondary damage remains undetected. Secondary damage may be produced in some structure, remote from the location of the primary damage, by the transmission of the damaging load along the tube. Damage of this nature usually occurs where the most abrupt change in direction of load travel is experienced. If this damage remains undetected, subsequent normal loads may cause failure of the part.

b. Location and Alignment of Welds. Unless otherwise noted, welded steel tubing may be spliced or repaired at any location along the length of the tube. To avoid distortion, pay particular attention to the proper fit and alignment.

c. Members Dented at a Cluster. Repair dents at a steel-tube cluster joint by welding a specially formed steel patch plate over the dented area and surrounding tubes. (See Figure 4-34.) To prepare the patch plate, cut a section of steel sheet of the same material and thickness as the heaviest tube damaged. Trim the reinforcement plate so that the fingers extend over the tubes a minimum of 1.5 times the respective tube diameter. (See Figure 4-34.) Remove all the existing finish on the damaged cluster-joint area to be covered by the reinforcement plate. The reinforcement plate may be formed before any welding is attempted, or it may be cut and tack-welded to one or more of the tubes in the cluster joint, then heated and formed around the joint to produce a smooth contour. Apply sufficient heat to the plate while forming so that there is generally a gap of no more than 1/16 inch from the contour of the joint to the plate. In this operation avoid unnecessary heating, and exercise care to prevent damage at the point of the angle formed by any two adjacent fingers of the plate. After the plate is formed and tack-welded to the cluster joint, weld all the plate edges to the cluster joint.

d. Members Dented in a Bay. Repair dented, bent, cracked, or otherwise damaged tubular members by using a split-sleeve reinforcement. Carefully straighten the damaged member, and in the case of cracks, drill No. 40 (0.098 inch) stop holes at the ends of the crack.

4-92. REPAIR BY WELDED SLEEVE

This repair is outlined in Figure 4-35. Select a length of steel tube sleeve having an inside diameter approximately equal to the outside diameter of the damaged tube and of the same material, and at least the same wall thickness. Diagonally cut the sleeve reinforcement at a 30-degree angle on both ends so that the minimum distance of the sleeve from the edge of the crack or dent is not less than 1-1/2 times the diameter of the damaged tube. Cut through the entire length of the reinforcement sleeve, and separate the half-sections of the sleeve. Clamp the two sleeve sections to the proper positions on the affected areas of the original tube. Weld the reinforcement sleeve along the length of the two sides, and weld both ends of the sleeve to the damaged tube. (See Figure 4-35.) The filling of dents or cracks with welding rod in lieu of reinforcing the member is not acceptable.

4-93. REPAIR BY BOLTED SLEEVE

Do not use bolted-sleeve repairs on welded steel-tube structure unless specifically authorized by the manufacturer or the FAA. The tube area removed by the bolt holes, in this type of repair, may prove critical.

4-94. WELDED-PATCH REPAIR

Dents or holes in tubing may be repaired by using a patch of the same material, one gauge thicker. (See Figure 4-36.)

b. Punctured Tubing. Holes are not longer than tube diameter and involve not more than 1/4 of tube circumference.

4-95. SPLICING TUBING BY INNER-SLEEVE METHOD

If the damage to a structural tube is such that a partial replacement of the tube is necessary, the inner-sleeve splice is recommended; especially where a smooth tube surface is desired. (See Figure 4-37.)

4-96. SPLICING TUBING BY OUTER-SLEEVE METHOD

If partial replacement of a tube is necessary, make the outer-sleeve splice using a replacement tube of the same diameter. Since the outer-sleeve splice requires the greatest amount of welding, it should be used only when the other splicing methods are not suitable. Information on the replacement by use of the outer-sleeve method is given in Figure 4-38 and Figure 4-39.

4-97. SPLICING USING LARGER DIAMETER REPLACEMENT TUBES

The method of splicing structural tubes, as shown in Figure 4-40, requires the least amount of cutting and welding. However, this splicing method cannot be used where the damaged tube is cut too near the adjacent cluster joints, or where bracket-mounting provisions make it necessary to maintain the same replacement tube diameter as the original. As an aid to installing the replacement tube, squarely cut the original damaged tube leaving a minimum short stub equal to 2-1/2-tube diameters on one end and a minimum long stub equal to 4-1/2-tube diameters on the other end. Select a length of steel tube of the same material and at least the same wall thickness, having an inside diameter approximately equal to the outside diameter of the damaged tube. Fit this replacement tube material snugly around the original tube with a maximum diameter difference of 1/16 inch. From this replacement tube material, cut a section of tubing diagonally (or fishmouth) of such a length that each end of the tube is a minimum distance of 1-1/2-tube diameters from the end of the cut on the original tube. Use a fishmouth cut replacement tube wherever possible. Deburr the edges of the replacement tube and original tube stubs. If a fishmouth cut is used, file out the sharp radius of the cut with a small round file. Spring the long stub of the original tube from the normal position, slip the replacement tube over the long stub, and then back over the short stub. Center the replacement tube between the stubs of the original tube. Tack weld one end of the replacement tube in several places, then weld completely around the end. In order to prevent distortion, allow the weld to cool completely, then weld the remaining end of the replacement tube to the original tube.

4-98. REPAIRS AT BUILT-IN FUSELAGE FITTINGS

Make splices in accordance with the methods described in paragraphs 4-86 through 4-92. Repair built-in fuselage fittings in the manner shown in Figure 4-41. The following paragraphs outline the different methods as shown in Figure 4-41.

a. Tube of Larger Diameter Than Original. A tube (sleeve) of larger diameter than the original is used in the method shown in Figure 4-41, image (A). The forward splice is a 30-degree scarf splice. Cut the rear longeron (right) approximately 4 inches from the centerline of the joint and fit a 1 inch long spacer over the longeron, and edge weld this spacer and longeron. Make a tapered “V” cut approximately 2 inches long in the aft end of the outer sleeve, and swage the end of the outer sleeve to fit the longeron and weld.

b. Tube of Same Diameter as Original. In the method shown in Figure 4-41, image (B), the new section is the same size as the longeron forward (left) of the fitting. The rear end (right) of the tube is cut at 30 degrees and forms the outside sleeve of the scarf splice. A sleeve is centered over the forward joint as indicated.

c. Simple Sleeve. In Figure 4-41, image (C), it is assumed the longeron is the same size on each side of the fitting. It is repaired by a sleeve of larger diameter than the longeron.

d. Large Difference in Longeron Diameter Each Side of Fitting. Figure 4-41, image (D) assumes that there is 1/4-inch difference in the diameter of the longeron on the two sides of the fitting. The section of longeron forward (left) of the fitting is cut at 30 degrees, and a section of tubing of the same size as the tube and of such length as to extend well to the rear (right) of the fitting is slipped through it. One end is cut at 30 degrees to fit the 30-degree scarf at left, and the other end fishmouthed. This makes it possible to insert a tube of proper diameter to form an inside sleeve for the tube on the left of the fitting and an outside sleeve for the tube on the right of the fitting.

4-99. ENGINE-MOUNT REPAIRS

All welding on an engine mount must be of the highest quality, since vibration tends to accentuate any minor defect. Engine-mount members should preferably be repaired by using a larger diameter replacement tube, telescoped over the stub of the original member, and using fishmouth and rosette welds. However, 30-degree scarf welds in place of the fishmouth welds will be considered acceptable for engine-mount repair work.

4-100. BUILT-UP TUBULAR WING OR TAIL-SPARS

Repair built-up tubular wing or tail-spars by using any of the applicable splices and methods of repair shown in Figure 4-35 through Figure 4-45, provided the spars are not heat treated. In the case of heat-treated spars, the entire spar assembly would have to be reheat treated to the manufacturer’s specifications after completion of the repair. In general, this will be found less practicable than replacing the spar with one furnished by the manufacturer or holder of the Parts Manufacturer Approval (PMA) for the part.

4-101. WING-BRACE STRUTS AND TAIL-BRACE STRUTS

In general, it will be found advantageous to replace damaged wing-brace struts made either from rounded or streamlined tubing with new members purchased from the original manufacturer. However, there is no objection, from an airworthiness point of view, to repairing such members in a proper manner. An acceptable method of repair, if streamlined tubing is used, will be found in Figure 4-43. Repair similar members made of round tubes using a standard splice, as shown in Figure 4-35, Figure 4-37, or Figure 4-38.

a. Location of Splices. Steel-brace struts may be spliced at any point along the length of the strut, provided the splice does not overlap part of an end fitting. The jury-strut attachment is not considered an end fitting; therefore, a splice may be made at this point. The repair procedure and workmanship minimize distortion due to welding and the necessity for subsequent straightening operations. Observe every repaired strut carefully during initial flights to ascertain that the vibration characteristics of the strut and attaching components are not adversely affected by the repair. A wide range of speed and engine-power combination must be covered during this check.

b. Fit and Alignment. When making repairs to wing and tail surface brace members, ensure repairs are properly fit and aligned to avoid distortion.

4-102. LANDING GEAR REPAIR

a. Round Tube Construction. Repair landing gears made of round tubing using standard repairs and splices as shown in Figure 4-35 and Figure 4-41.

b. Streamline Tube Construction. Repair landing gears made of streamlined tubing by either one of the methods shown in Figure 4-42, Figure 4-44, or Figure 4-45.

c. Axle Assemblies. Representative types of repairable and nonrepairable landing gear axle assemblies are shown in Figures 4-46 and 4-47. The types as shown in Figure 4-46, images (A), (B), and (C) are formed from steel tubing and may be repaired by the applicable method shown in Figure 4-35 through Figure 4-45. However, it will always be necessary to ascertain whether or not the members are heat treated. The axle assembly as shown in Figure 4-47 is, in general, of a nonrepairable type for the following reasons.

4-103. REPAIRS TO WELDED ASSEMBLIES

These repairs may be made by the following methods.

4-104. STAINLESS STEEL STRUCTURE

Repair structural components made from stainless steel, particularly the “18-8” variety (18 percent chromium, 8 percent nickel), joined by spot welding, in accordance with the instructions furnished by the manufacturer, a Designated Engineering Representative (DER), or the FAA. Substitution of bolted or riveted connections for spot-welded joints are to be specifically approved by a DER or the FAA. Repair secondary structural and nonstructural elements such as tip bows or leading and trailing edge tip strips of wing and control surfaces by soldering with a 50-50 lead-tin solder or a 60-40 lead-tin solder. For best results, use a flux of phosphoric acid (syrup). Since the purpose of flux is to attack the metal so that the soldering will be effective, remove excess flux by washing the joint. Due to the high heat conductivity of the stainless steel, use a soldering iron large enough to do the work properly.

4-105.–4-110. [RESERVED.]

SECTION 6. WELDING AND BRAZING SAFETY

4-111. GENERAL

A number of inherent hazards exist in the use of oxy-fuel welding and cutting apparatus. It is necessary that proper safety and operating procedures are understood. A thorough understanding of the proper safety and operating procedures minimizes the hazards involved and adds to the pleasure and efficiency of your work.

4-112. FIRE AND EXPLOSION SAFETY

Fires occur in welding areas because flammables are left where they can be ignited by welding sparks or gas welding flames. Before welding, clear the welding area of all flammables such as rags, paper, wood, paint cans, solvent, and trash containers. Do not weld in areas where flammables are present.

4-113. WELDING WORK AREA

CAUTION: Oil and grease in the presence of oxygen can ignite and burn violently.

4-114. FIRE PROTECTION

Practice fire prevention techniques whenever oxy-fuel operations are in progress. Simple precautions prevent most fires, and minimize damage in the event that a fire does occur. Always practice the following rules and safety procedures.

j. NEVER TEST FOR GAS LEAKS WITH A FLAME. Use an approved leak-detector solution.

4-115. PROTECTIVE APPAREL

b. Wear protective gloves, sleeves, aprons, and lace-up shoes to protect skin and clothing from sparks and slag.

CAUTION: Keep all clothing and protective apparel absolutely free of oil or grease.

4-116. FIRST AID KITS

Always keep a special welder’s first aid kit where it is easily accessible. Burns are the most common welding accidents.

4-117.–4-128. [RESERVED.]