Par 4-23 Page 4-11
b. Test for distinguishing heat-treatable and nonheat-treatable aluminum alloys. If for any reason the identification mark of the alloy is not on the material, it is possible to distinguish between some heat-treatable alloys and some nonheat-treatable alloys by immersing a sample of the material in a 10 percent solution of caustic soda (sodium hydroxide). Those heat-treated alloys containing several percent of copper (2014, 2017, and 2024) will turn black due to the copper content. High-copper alloys when clad will not turn black on the surface, but the edges will turn black at the center of the sheet where the core is exposed. If the alloy does not turn black in the caustic soda solution it is not evidence that the alloy is nonheat-treatable, as various high-strength heat-treatable alloys are not based primarily on the use of copper as an alloying agent. These include, among others, 6053, 6061, and 7075 alloys. The composition and heat-treating ability of alloys which do not turn black in a caustic soda solution can be established only by chemical or spectro-analysis.
4-24.–4-35. [RESERVED.]
Par 4-36 Page 4-12 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.
(1) Free play and stiffness may best be measured by a simple static test where “upward” and “downward” (or “leftward” and “rightward”) point forces are applied near the trailing edge of the tab at the span-wise attachment of the actuator (so as not to twist the tab). The control surface to which the trim tab is attached should be locked in place. Rotational deflection readings are then taken near the tab trailing edge using an appropriate measuring device, such as a dial gauge. Several deflection readings should be taken using loads first applied in one direction, then in the opposite. If the tab span does not exceed 35 percent of the span of the supporting control surface, the total free play at the tab trailing edge should not exceed 2 percent of the tab chord. If the tab span equals or exceeds 35 percent of the span of the supporting control surface, the total free play at the tab trailing edge should not exceed 1 percent of the distance from the tab hinge line to the trailing edge of the tab perpendicular to the tab hinge line. For example, a tab that has a chord of 4 inches and less than or equal to 35 percent of the control surface span would have a
Par 4-36 Page 4-13 maximum permissible free play of 4 inches x 0.020 or 0.080 inches (total motion up and down) measured at the trailing edge. Correct any free play in excess of this amount.
(2) Care must also be exercised during repair or rework to prevent stress concentration points or areas that could increase the fatigue susceptibility of the trim tab system. Advisory Circular (AC) 23.629-1, Means of Compliance with Title 14 CFR, Part 23, § 23.629, Flutter, contains additional information on this subject.
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.
a. Six types of major stresses are known and should be considered when making repairs.
These are tension, compression, bending, torsion, shear, and bearing.
b. The design of an aircraft repair is complicated by the requirement that it be as light as possible. If weight were not critical, repairs could be made with a large margin of safety. But in actual practice, repairs must be strong enough to carry all of the loads with the required factor of safety, but they must not have too much extra strength. A joint that is too weak cannot be tolerated, but neither can one that is too strong because it can create stress risers that may cause cracks in other locations.
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.]
Par 4-50 Page 4-14 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.
a. Aircraft Principal Structural Elements (PSE) and joints are designed to carry loads by distributing them as stresses. The elements and joints as originally fabricated are strong enough to resist these stresses, and must remain so after any repairs. Long, thin elements are called members. Some examples of members are the metal tubes that form engine mount and fuselage trusses and frames, beams used as wing spars, and longerons and stringers of metal-skinned fuselages and wings. Longerons and stringers are designed to carry principally axial loads, but are sometimes required to carry side loads and bending moments, as when they frame cutouts in metal-skinned structures. Truss members are designed to carry axial (tension and compression) loads applied to their ends only. Frame members are designed to carry side loads and bending moments in addition to axial loads. Beam members are designed to carry side loads and bending moments that are usually large compared to their axial loads. Beams that must resist large axial loads, particularly compression loads, in combination with side loads and bending moments are called beam-columns. Other structural elements such as metal skins, plates, shells, wing ribs, bulkheads, ring frames, intercostal members, gussets, other reinforcements, and fittings are designed to resist complex stresses, sometimes in three dimensions.
b. Any repair made on an aircraft structure must allow all of the stresses to enter, sustain these stresses, and then allow them to return into the structure. The repair must be equal to the original structure, but not stronger or stiffer, which will cause stress concentrations or alter the resonant frequency of the structure.
c. All-metal aircraft are made of very thin sheet metal, and it is possible to restore the strength of the skin without restoring its rigidity. All repairs should be made using the same type and thickness of material that was used in the original structure. If the original skin had corrugations or flanges for rigidity, these must be preserved and strengthened. If a flange or corrugation is dented or cracked, the material loses much of its rigidity; and it must be repaired in such a way that will restore its rigidity, stiffness, and strength.
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
Par 4-52 Page 4-15 diameter of the rivet. Therefore, it is important that the proper drill size be used for each rivet diameter.
(1) The acceptable drill size for rivets may be found in MIL-HDBK-5, Metallic Materials and Elements for Aerospace Vehicle Structures, or MMPDS-10.
(2) Avoid drilling oversized holes or otherwise decreasing the effective tensile areas of wing-spar cap strips, wing, fuselage, fin longitudinal stringers, or highly stressed tensile members. Make all repairs, or reinforcements, to such members in accordance with factory recommendations or with the specific approval of an FAA representative.
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.
Par 4-54 Page 4-16
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.
a. The minimum radius is determined by the composition of the metal, its temper, and thickness. Table 4-6 shows the recommended radius for different types of aluminum. Note that the smaller the thickness of the material, the smaller the recommended minimum BR, and that as the material increases in hardness, the recommended BR increases.
b. When using layout techniques, the mechanic must be able to calculate exactly how much material will be required for the bend. It is easier to lay out the part on a flat sheet before the bending or shaping is performed. Before bending, smooth all rough edges, remove burrs, and drill relief holes at the ends of bend lines and at corners to prevent cracks from starting. Bend lines should preferably be made to lie at an angle to the grain of the metal (preferably 90 degrees).
c. BR in inches for a specific metal composition (alloy) and temper is determined from 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.
a. Setback is a measurement used in sheet metal layout. It is the distance the jaws of a brake must be setback from the mold line to form a bend. For a 90-degree bend, the point is back from the mold line to a distance equal to the BR plus the metal thickness. The mold line is an extension of the flat side of a part beyond the radius. The mold line dimension of a part is the dimension made to the intersection of mold lines, and is the dimension the part would have if its corners had no radius. (See Figure 4-2.)
Par 4-56 Page 4-17 TABLE 4-6. RECOMMENDED RADII FOR 90-DEGREE BENDS IN
ALUMINUM ALLOYS Alloy and Temper Approximate Sheet Thickness (t) (inch)
0.016 0.032 0.064 0.128 0.182 0.258 2024-01 0 0–1t 0–1t 0–1t 0–1t 0–1t 2024-T31, 2 1½t–3t 2t–4t 3t–5t 4t–6t 4t–6t 5t–7t 2024-T61 2t–4t 3t–5t 3t–5t 4t–6t 5t–7t 6t–10t 5052-0 0 0 0–1t 0–1t 0–1t 0–1t 5052-H32 0 0 ½t–1t ½t–1½t ½t–1½t ½t–1½t 5052-H34 0 0 ½t–1½t 1½t–2½t 1½t–2½t 2t–3t 5052-H36 0–1t ½t–1½t 1t–2t 1½t–3t 2t–4t 2t–4t 5052-H38 ½t–1½t 1t–2t 1½t–3t 2t–4t 3t–5t 4t–6t 6061-0 0 0–1t 0–1t 0–1t 0–1t 0–1t 6061-T4 0–1t 0–1t ½t–1½t 1t–2t 1½t–3t 2½t–4t 6061-T6 0–1t ½t–1½t 1t–2t 1½t–3t 2t–4t 3t–4t 7075-0 0 0–1t 0–1t ½t–1½t 1t–2t 1½t–3t 7075-T61 2t–4t 3t–5t 4t–6t 5t–7t 5t–7t 6t–10t 1 Alclad sheet may be bent over slightly smaller radii than the corresponding tempers of uncoated alloy.
2 Immediately after quenching, this alloy may be formed over appreciably smaller radii.
FIGURE 4-2. SETBACK FOR A 90-DEGREE BEND
Par 4-56 Page 4-18
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.
TABLE 4-7. K-CHART FOR DETERMINING SETBACK FOR BENDS OTHER THAN 90 DEGREES Deg. K Deg. K Deg. K Deg. K Deg. K 1 0.0087 37 0.3346 73 0.7399 109 1.401 145 3.171 2 0.0174 38 0.3443 74 0.7535 110 1,428 146 3.270 3 0.0261 39 0.3541 75 0.7673 111 1.455 147 3.375 4 0.0349 40 0.3639 76 0.7812 112 1.482 148 3.487 5 0.0436 41 0.3738 77 0.7954 113 1.510 149 3.605 6 0.0524 42 0.3838 78 0.8097 114 1.539 150 3.732 7 0.0611 43 0.3939 79 0.8243 115 1.569 151 3.866 8 0.0699 44 0.4040 80 0.8391 116 1.600 152 4.010 9 0.0787 45 0.4142 81 0.8540 117 1.631 153 4.165 10 0.0874 46 0.4244 82 0.8692 118 1.664 154 4.331 11 0.0963 47 0.4348 83 0.8847 119 1.697 155 4.510 12 0.1051 48 0.4452 84 0.9004 120 1.732 156 4.704 13 0.1139 49 0.4557 85 0.9163 121 1.767 157 4.915 14 0.1228 50 0.4663 86 0.9324 122 1.804 158 5.144 15 0.1316 51 0.4769 87 0.9489 123 1.841 159 5.399 16 0.1405 52 0.4877 88 0.9656 124 1.880 160 5.671 17 0.1494 53 0.4985 89 0.9827 125 1.921 161 5.975 18 0.1583 54 0.5095 90 1.000 126 1.962 162 6.313 19 0.1673 55 0.5205 91 1.017 127 2.005 163 6.691 20 0.1763 56 0.5317 92 1.035 128 2.050 164 7.115 21 0.1853 57 0.5429 93 1.053 129 2.096 165 7.595 22 0.1943 58 0.5543 94 1.072 130 2.144 166 8.144 23 0.2034 59 0.5657 95 1.091 131 2.194 167 8.776 24 0.2125 60 0.5773 96 1.110 132 2.246 168 9.514 25 0.2216 61 0.5890 97 1.130 133 2.299 169 10.38 26 0.2308 62 0.6008 98 1.150 134 2.355 170 11.43 27 0.2400 63 0.6128 99 1.170 135 2.414 171 12.70 28 0.2493 64 0.6248 100 1.191 136 2.475 172 14.30 29 0.2586 65 0.6370 101 1.213 137 2.538 173 16.35 30 0.2679 66 0.6494 102 1.234 138 2.605 174 19.08 31 0.2773 67 0.6618 103 1.257 139 2.674 175 22.90 32 0.2867 68 0.6745 104 1.279 140 2.747 176 26.63 33 0.2962 69 0.6872 105 1.303 141 2.823 177 38.18 34 0.3057 70 0.7002 106 1.327 142 2.904 178 57.29 35 0.3153 71 0.7132 107 1.351 143 2.988 179 114.59 36 0.3249 72 0.7265 108 1.376 144 3.077 180 Inf.
(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:
Setback = K(BR + MT) = 0.41421 (0.25 + 0.064) = 0.130 inches FIGURE 4-3. METHODS OF DETERMINING SETBACK FOR BENDS OTHER THAN 90 DEGREES
• T = 0.064"
- R = 0.250
SETBACK
SETBACK This is a 90-degree angle.
Setback = BR + MT = 0.250 + 0.064 = 0.314" T = 0.064" R = 0.250
SETBACK This is a 135-degree open angle. The metal has only been bent 45°.
SB = (BR + MT)K = (0.250 + 0.064) 0.4121 = 0.130" T = 0.064 -
SETBACK
- R = 0.250
SETBACK This is a 45-degree closed angle.
Setback = (BR + MT)K = (0.250 + 0.064)2.4142 = 0.758"
(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.
a. The two major types of rivets used in aircraft are the common solid shank rivet, which must be driven using an air-driven rivet gun and bucking bar; and special (blind) rivets, which Par 4-56 Page 4-19
Par 4-57 Page 4-20 are installed with special installation tools. Design allowables for riveted assemblies are specified in MIL-HDBK-5 or MMPDS-10.
(1) Solid shank rivets are used widely during assembly and repair work. They are identified by the material of which they are made, the head type, size of shank, and temper condition.
(2) The material used for the majority of solid shank rivets is aluminum alloy. The strength and temper conditions of aluminum alloy rivets are identified by digits and letters similar to those used to identify sheet stock. The 1100, 2017-T, 2024-T, 2117-T, and 5056 rivets are the six grades usually available. Army/Navy (AN)-type aircraft solid rivets can be identified by code markings on the rivet heads. A rivet made of 1100 material is designated as an “A” rivet, and has no head marking. The 2017-T alloy rivet is designated as a “D” rivet and has a raised teat on the head. Two dashes on a rivet head indicate a 2024-T alloy designated as a “DD” rivet.
The 2117-T rivet is designated as an “AD” rivet, and has a dimple on the head. A “B” designation is given to a rivet of 5056 material and is marked with a raised cross on the rivet head. Each type of rivet is identified by a part number to allow the user to select the correct rivet.
The numbers are in series and each series represents a particular type of head. (See Figure 4-4 and Table 4-8.)
FIGURE 4-4. RIVET IDENTIFICATION AND PART NUMBER BREAKDOWN
RIVET IDENTIFICATION The material can be identified by the head marking Material Rivet Code Head Marking Material O¢ A PLAIN (Dyed) 1100 AD DIMPLED 2117 D RAISED DOT 2017T DO TWO RAISED DASHES 2024 RAISED CROSS (Dyed) 5056 RAISED CIRCLE 7050 M TWO DOTS Monel
PART NUMBER BREAKDOWN
• MS20470 AD 4-7
•• MS20470 • AD Type Head Material Diameter Length Code Measured Measured in 32nds in 16ths of an inch of an inch Length is measured from the top of the flush head and the underside of the universal head.
Par 4-57 Page 4-21
TABLE 4-8. AIRCRAFT RIVET IDENTIFICATION Material 1100 2117T 20171 2017T-HD 2024T 5056T 7075-T73 Head Plain Dimpled Raised Raised Raised Raised Three Marking Dot Dot Double Cross Raised Dash Dashes AN
AD D D DD Material Code AN425 78• Counter- Sunk Head AN426 100• Counter- Sunk Head MS20426 AN427 100• Counter- Sunk Head MS20427 AN430 Round Head MS20470 AN435 Round Head MS20613 MS20615 AN 441 Flat Head AN 442 Flat Head MS20470 AN 455 Brazier × Head MS20470 AN 456 Brazier Head MS20470 AN 470 Universal Head MS20470 Heat Treat Before No No Yes No Yes No No Using Shear Strength 10000 30000 34000 38000 41000 27000 psi Bearing Strength 25000 100000 113000 126000 136000 90000 psi Par 4-57 Page 4-22