Par 4-1 Page 4-103
CHAPTER 4. METAL STRUCTURE, WELDING, AND BRAZING
SECTION 1. IDENTIFICATION OF METALS
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.
Ferrous (iron) alloy materials are generally classified according to carbon content. (See table 4-1.)
TABLE 4-1. Ferrous (iron) alloy materials.
MATERIALS CARBON CONTENT Wrought iron Trace to 0.08% Low carbon steel 0.08% to 0.30% Medium carbon steel 0.30% to 0.60% High carbon steel 0.60% to 2.2% Cast iron 2.3% to 4.5%
The strength and ductility, or toughness of steel, is controlled by the kind and quantity of alloys used and also by cold-working or heat-treating processes used in manufacturing. In general, any process that increases the strength of a material will also decrease its ductility.
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.
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 chromiummolybdenum 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.
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.
“1025” welded tubing as per Specification MIL-T-5066 and “1025” seamless tubing conforming to Specification MIL-T-5066A are interchangeable.
Par 4-1 Page 4-104
INTERCHANGEABILITY OF STEEL TUBING .
“4130” welded tubing conforming to Specification MIL-T-6731, and “4130” seam-less tubing conforming to Specification MIL-T-6736 are interchangeable.
NE-8630 welded tubing conforming to Specification MIL-T-6734, and NE-8630 seamless tubing conforming to Specification MIL-T-6732 are interchangeable.
IDENTIFICATION OF ALUMINUM. To provide a visual means for identifying the various grades of aluminum and aluminum alloys, such metals are usually marked with symbols such as a Government Specification Number, the temper or condition furnished, or the commercial code marking. Plate and sheet are usually marked with specification numbers or code markings in rows approximately 5 inches apart. Tubes, bars, rods, and extruded shapes are marked with specification numbers or code markings at intervals of 3 to 5 feet along the length of each piece.
The commercial code marking consists of a number which identifies the particular composition of the alloy. In addition, letter suffixes (see table 4-4) designate the basic temper designations and subdivisions of aluminum alloys.
TABLE 4-2. Numerical system for steel identification.
TYPES OF STEELS NUMERALS
AND DIGITS Plain carbon steel 10XX Carbon steel with additional sulfur for easy machining.
11XX Carbon steel with about 1.75% manganese 13XX .25% molybdenum. 40XX 1% chromium, .25% molybdenum 41XX 2% nickel, 1% chromium, .25% molybdenum 43XX
1.7% nickel, .2% molybdenum 46XX
3.5% nickel, .25% molybdenum 48XX 1% chromium steels 51XX 1% chromium, 1.00% carbon 51XXX
1.5% chromium steels 52XX
1.5% chromium, 1.00% carbon 52XXX 1% chromium steel with .15% vanadium 61XX .5% chromium, .5% nickel, .20% molybdenum 86XX .5% chromium, .5% nickel, .25% molybdenum 87XX 2% silicon steels, .85% manganese 92XX
3.25% nickel, 1.20% chromium, .12% molybdenum 93XX
TABLE 4-3. Examples of stainless and heat-resistant steels nominal composition (percent)
ALLOY DESIGNATION CARBON CHROMIUM NICKEL OTHER GENERAL CLASS OF STEEL 302 0.15 18 9 Austenitic 310 0.25 25 20 Austenitic 321 0.08 18 11 Titanium Austenitic 347 0.08 18 11 Columbium or Tantalum Austenitic 410 0.15 12.5 Martensitic, Magnetic 430 0.12 17 Ferritic, Magnetic 446 0.20 25 Nitrogen Ferritic, Magnetic PH15-7 Mo 0.09 15 7 Molybdenum, Aluminum Precipitation Hardening 17-4 PH 0.07 16.5 4 Copper, Columbium or Tantalum Precipitation Hardening
Par 4-4 Page 4-105 (and 4-4) TABLE 4-4. Basic temper designations and subdivisions from aluminum alloys.
NON HEAT-TREATABLE ALLOYS HEAT-TREATABLE ALLOYS Temper Designation
Definition Temper Designation
Definition -0 Annealed recrystallized (wrought products only) applies to softest temper of wrought products.
-0 Annealed recrystallized (wrought products only) applies to softest temper of wrought products.
-H1 Strain-hardened only. Applies to products which are strain-hardened to obtain the desired strength without supplementary thermal treatment.
-T1 Cooled from an elevated temperature shaping process (such as extrusion or casting) and naturally aged to a substantially stable condition.
-H12 Strain-hardened one-quarter-hard tem per. -T2 Annealed (castings only).
-H14 Strain-hardened half-hard temper. -T3 Solution heat-treated and cold-worked by the flattening or straightening operation.
-H16 Strain-hardened three-quarters-hard temper. -T36 Solution heat-treated and cold-worked by reduction of 6 percent -H18 Strain-hardened full-hard temper. -T4 Solution heat-treated.
-H2 Strain-hardened and then partially annealed.
Applies to products which are strain-hardened more than the desired final amount and then reduced in strength to the desired level by partial annealing.
-T42 Solution heat-treated by the user regardless of prior temper (applicable only to 2014 and 2024 alloys).
-H22 Strain-hardened and partially annealed to one-quarter-hard temper.
-T5 Artificially aged only (castings only).
-H24 Strain-hardened and partially annealed to half-hard temper.
-T6 Solution heat-treated and artificially aged.
-H26 Strain-hardened and partially annealed to three-quarters-hard temper.
-T62 Solution heat-treated and aged by user regardless of prior temper (applicable only to 2014 and 2024 alloys).
-H28 Strain-hardened and partially annealed to full-hard temper.
-T351, -T451, -T3510, -T3511, -T4510, -T4511.
Solution heat-treated and stress relieved by stretching to produce a permanent set of 1 to 3 percent, depending on the product.
-H3 Strain-hardened and then stabilized. Applies to products which are strain-hardened and then stabilized by a low temperature heating to slightly lower their strength and increase ductility.
-T651, -T851, -T6510, -T8510, -T6511, -T8511.
Solution heat-treated, stress relieved by stretching to produce a permanent set of 1 to 3 percent, and artificially aged.
-H32 Strain-hardened and then stabilized. Final temper is one-quarter hard.
-T652 Solution heat-treated, compressed to produce a permanent set and then artificially aged.
-H34 Strain-hardened and then stabilized. Final temper is one-half hard.
-T8 Solution heat-treated, cold-worked and then artificially aged.
-H36 Strain-hardened and then stabilized. Final temper is three-quarters hard.
-T/4 Solution heat-treated, cold-worked by the flattening or straightening operation, and then artificially aged.
-H38 Strain-hardened and then stabilized. Final temper is full-hard.
-T86 Solution heat-treated, cold-worked by reduction of 6 percent, and then artificially aged.
-H112 As fabricated; with specified mechanical property limits.
-T9 Solution heat-treated, artificially aged and then cold-worked.
-F For wrought alloys; as fabricated. No mechanical properties limits. For cast alloys; as cast.
-T10 Cooled from an elevated temperature shaping process artificially aged and then cold-worked.
-F For wrought alloys; as fabricated. No mechanical properties limits. For cast alloys; as cast.
4-5.4-15. [RESERVED.]
Par 4-16 Page 4-106
SECTION 2. TESTING OF METALS
HARDNESS TESTING. If the material type is known, a hardness test is a simple way to verify that the part has been properly heattreated. 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, under-strength 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 heattreated 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.
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 American Society of Testing Materials (ASTM E-18) specifications.
BRINELL HARDNESS TEST. In this test a standard constant load, usually 500 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 is the load, in kg; D is the diameter of the ball, in mm; and d is the diameter of the indentation, in mm.
Par 4-20 Page 4-107 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.
Do not make indentations on a curved surface having a radius of less than 1 inch.
Do make the indentations with the correct spacing. Indentations should not be made too close to the edge of the work piece being tested.
Apply the load steadily to avoid overloading caused by inertia of the weights.
Apply the load so the direction of loading and the test surface are perpendicular to each other within 2 degrees.
The thickness of the work piece being tested should be such that no bulge or mark showing the effect of the load appears on the side of the work piece opposite the indentation.
The indentation diameter should be clearly outlined.
Limitations. The Brinell hardness test has three principal limitations.
The work piece must be capable of accommodating the relatively large indentations.
Due to the relatively large indentations, the work piece should not be used after testing.
The limit of hardness, 15 HB with the 500-kg load to 627 HB with the 3,000-kg load, is generally considered the practical range.
Calibration. A Brinell Hardness Tester should be calibrated to meet ASTM standard E10 specifications.
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.
The indenter is made of diamond, and is in the form of a square-based pyramid having an angle of 136 degrees between faces. The facets are highly-polished, free from surface imperfections, and the point is sharp. The loads applied vary from 1 to 120 kg; the standard loads are 5, 10, 20, 30, 50, 100, and 120 kg. For most hardness testing, 50 kg is maximum.
A Vickers hardness tester should be calibrated to meet ASTM standard E10 specifi cations, acceptable for use over a loading range.
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 to 500 g. In general, the term is related to the size of the indentation rather than to the load applied.
Par 4-20 Page 4-108 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 Br inell or Rockwell, and are as follows.
Measuring the hardness of precision work pieces that are too small to be measured by the more common hardness-testing methods.
Measuring the hardness of product forms such as foil or wire that are too thin or too small in diameter to be measured by the more conventional methods.
Monitoring of carburizing or nitriding operations, which is sometimes accomplished by hardness surveys taken on cross sections of test pieces that accompanied the work pieces through production operations.
Measuring the hardness of individual microconstituents.
Measuring the hardness close to edges, thus detecting undesirable surface conditions such as grinding burn and decarburization.
Measuring the hardness of surface layers such as plating or bonded layers.
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 the more widely used in Europe.
Knoop indentation testing is performed with a diamond, ground to pyramidal form, that produces a diamond-shaped indentation with an approximate ratio between long and short diagonals of 7 to 1. The indentation depth is about one-thirtieth of its length. Due to the shape of the indenter, indentations of accurately measurable length are obtained with light loads.
The Knoop hardness number (HK) is the ratio of the load applied to the indenter to the unrecovered projected area of indentation.
The formula for this follows.
H K=P/A =P/ Cl2
Where P is the applied load, in kg; A is the unrecovered projected area of indentation, in square mm; l is the measured length of the long diagonal, in mm; and C is 0.07028, a constant of the indenter relating projected area of the indentation to the square of the length of the long diagonal.
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.)
FIGURE 4-1. Comparison of indentation made by Knoop and Vickers indenters in the same work metal and at the same loads.
Par 4-36 Page 4-109 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.
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.
Clad aluminum alloys have surface layers of pure aluminum or corro sion-resistant aluminum alloy bonded to the core material to inhibit corrosion. Presence of such a coating may be determined under a magnifying glass by examination of the edge surface which will show three distinct layers. In aluminum alloys, the properties of an y specific alloy can be altered by work hardening (often called strain-hardening), heat treatment, or by a combination of these processes.
Test for distinguishing heat-treatabl e 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
Par 4-36 Page 4-110 and some nonheat-treatable alloys by immersing a sample of the material in a 10 percent solution of caustic soda (sodium hydroxide). Those h eat-treated alloys containing several per cent 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 e xposed. 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.
TABLE 4-5. Hardness values for aluminum alloys. (Reference MIL-H-6088G .)
Material
Hardness
Brinell number Commercial Temper 500 kg. load Designation 10 mm. ball
1100 0 23 H18 44 3003 0 28 H16 47 2014 0 45 T6 135 2017 0 45 T6 105 2024 0 47 T4 120 2025 T6 110 6151 T6 100 5052 0 47 H36 73 6061 0 30 T4 65 T6 95 7075 T6 135 7079 T6 135 195 T6 75 220 T4 75 C355 T6 80 A356 T6 70
4-24.4-35. [RESERVED.]
Par 4-36 Page 4-111 SECTION 3. PRECAUTIONARY MEASURES
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.
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 fl utter 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 w ith the manufacturer’s instructions. Flight tes ting must be accomplished to ensure flutter is not a problem. Failure to check and retain control surface balan ce within the original or maximum allowable value could result in a serious flight hazard.
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 instru ctions relative to finishing and balance of control surfaces.
Trapped Water or Ice. Instances of flutter have occurred from unbalanced conditions caused by the collection of water or ice with in the surface. Therefore, ventilation and drainage provisions must be checked and retained when maintenance is being done.
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.
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
Par 4-50 Page 4-112 4 inches and less than or equal to 35 percent of the cont rol surface span would have a maximum permissible free play of 4 inches x 0.020 or 0.080 inches (total motion up and down) measured at the tr ailing edge. Correct any free play in excess of this amount.
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) 23.629-1A, Means of Compliance with Section 23.629, “Flutter,” contains additional information on this subject.
NOTE: If the pilot has experi enced flutter, or thinks he/she has, 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.
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.
Six types of major stresses are known and should be considered when making repairs. These are tension, compression, bending, torsion, shear, and bearing
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 safe ty. 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.
Par 4-50 Page 4-113 TRANSFER OF STRESSES WITH-IN 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 crosssectional 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.
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.
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. B ecause 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 4-114 SECTION 4. METAL REPAIR PROCEDURES
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.
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, and other reinforcements, and fittings are designed to resist complex stresses, sometimes in three dimensions.
Any repair made on an aircraft structure must allow all of the stresses to enter, sust ain 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.
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 ma terial loses much of its rigidity; and it must be repaired in such a way that will restore its rigidity, stiffness, and strength.
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 text books on metal structures. Methods for repair of the major structural members must be specifically approved by the Federal Aviation Administration (FAA).
ALUMINUM ALLOY STRUCTURES. Extensive repairs to damaged stressed skin on monocoque-types of aluminum alloy structures should preferably be made in accordance with specific manufacturer’s instructions or other FAA-approved source.
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.