Forward Fuselage Ncdj Drilling System

Interfacial and Over Coatings

Slurry Infiltration with Si Particulates

Preform Fabrication

Reaction Bonded SiC CMC

High Temperature Liquid Silicon Infiltration

Removal of Slurry Carrier Liquid

Fig. 10.20. Liquid Silicon Infiltration Process

Reaction Bonded SiC CMC

High Temperature Liquid Silicon Infiltration

Removal of Slurry Carrier Liquid

Fig. 10.20. Liquid Silicon Infiltration Process an interfacial coating, along with a SiC over-coating, to protect them from the liquid silicon. Before infiltration, a fine grained silicon carbide particulate is slurry cast into the fiber preform. After removal of the slurry carrier liquid, melt infiltration is usually done at 2550° F, or higher, and is usually complete within a few hours, as illustrated in Fig. 10.20. The liquid silicon bonds the silicon carbide particulates together and forms a matrix that is somewhat stronger and denser than that obtained by CVI. Since the resulting matrix can contain up to 50% unreacted silicon, the long-term use temperature is limited to about 2200° F.11 The amount of unreacted silicon can be reduced by infiltrating the preform with carbon slurries prior to the silicon infiltration process. However, some unreacted or free silicon will always be present in the matrix.

In another variant of the liquid metal infiltration process, liquid silicon is reacted with unprotected carbon fibers to form a silicon carbide matrix. After the initial step of using prepreg to form a highly porous carbon matrix part, liquid silicon is infiltrated into the structure, where it reacts with the carbon to form silicon carbide along with unreacted silicon and carbon. Since the carbon fibers are intended to react with the liquid silicon, no coatings are used on the fibers. However, the poor oxidation resistance of the carbon fibers means that the entire part will require an oxidation resistant coating.

The liquid metal infiltration processes have several advantages: (1) they produce a fairly dense SiC based matrix with a minimum of porosity; (2) the processing time is shorter than for most ceramic matrix composite fabrication processes; and (3) the dense and closed porosity on the surface can often eliminate the need for a final oxidation resistant coating. The major disadvantage is the high temperatures (>2550° F) required for liquid silicon infiltration that exposes the fibers to possible degradation, due to the high temperatures and corrosive nature of liquid silicon. In addition, the temperatures can be even higher due to the possibility of an exothermic reaction between silicon and carbon.

Summary

Monolithic high performance ceramics combine desirable characteristics, such as high strength and hardness, high temperature capability, chemical inertness, wear resistance, and low density. The greatest drawback of ceramics is their extremely low fracture toughness, which in practice means that these materials have a very low tolerance of crack-like defects. Incorporation of fibers, whiskers, or particles in a ceramic matrix can result in a tougher ceramic material, because the reinforcements introduce energy dissipating mechanisms, such as debonding at the fiber-to-matrix interface, crack deflection, fiber bridging, and fiber pullout. For these energy dissipating mechanisms to be effective, there must be a poor bond at the fiber-to-matrix interface.

Similar to MMCs, there are very few commercial applications for CMCs, due to their high costs and concerns for reliability; however, carbon-carbon has found applications in aerospace for thermal protection systems. Carboncarbon is capable of withstanding temperatures in excess of 3000° F provided it is protected against oxidation. Both internal and external oxidation protection systems are used, with SiC and glass forming sealer compounds being the most prevalent. C-C can be made by either pyrolysis of organic compounds or CVI. Both procedures are complex and lengthy, resulting in long lead times and high costs.

The slurry infiltration process is the most important technique used to produce glass and glass-ceramic composites. The slurry infiltration process followed by hot pressing is well suited for glass or glass-ceramic matrix composites, mainly because the processing temperatures for these materials are lower than those used for crystalline ceramics. For discontinuous CMCs, short fibers, whiskers, or particles are mixed with a ceramic powder slurry, dried, and hot pressed. For continuous fiber crystalline CMCs, either prepregs or preforms are used to maintain the fiber architecture. To provide protection to the fibers during processing and service and to insure weak interface bonding, interfacial coatings of C and BN are applied by CVD.

The most prevalent fabrication approaches for continuous fiber crystalline ceramic matrix composites include: PIP, CVI, DMO, and LSI.

The PIP process consists of (1) infiltration of the preform with the polymer; (2) consolidation of the impregnated preform; (3) cure of the polymer matrix to prevent melting during subsequent processing; and (4) pyrolysis of the cured polymer to convert it to a ceramic matrix. A number of re-infiltration and pyrolysis cycles, often 5-10, are required to produce a high density part. The biggest advantage of the polymer infiltration and pyrolysis process is the use of the familiar methods employed in organic matrix composite fabrication. However, the multiple infiltration and pyrolysis cycles are expensive and the lead times are long. In addition, it is almost impossible to fill all of the fine matrix cracks which degrade the mechanical properties and thermal conductivity.

In CVI, a solid is deposited within the open volume of a porous structure by the reaction or decomposition of gases or vapors. A porous preform of fibers is prepared and placed in a high temperature furnace. Reactant gases or vapors are then pumped into the chamber and flow around and diffuse into the preform. The gases decompose, or react, to deposit a solid onto and around the fibers. As the reaction progresses, the apparent diameter of the fibers increases and eventually fills the available porosity. The CVI method offers several advantages: (1) it is conducted at relatively low temperatures so damage to the fibers is minimal; (2) since most interfacial coatings are applied using CVD, the matrix infiltration can be conducted immediately after the interfacial coatings are applied; (3) it can also be used to fabricate fairly large and complex near net shapes; and (4) the mechanical and thermal properties are good because high purity matrices with controlled microstructures can be obtained. The major disadvantage of the CVI process is that it is not possible to obtain a fully dense part since the amount of residual porosity is around 10-15%, which adversely affects the mechanical and thermal properties. The other big disadvantage is that long processing times, often greater than 100 h, and multiple machining cycles result in high costs.

Directed metal oxidation, or reactive melt infiltration, uses liquid aluminum that reacts with air (oxygen) to form alumina (Al203), or with nitrogen to form aluminum nitride (AlN). This process is relatively low cost with near net shape capabilities, and complex shaped parts can be fabricated. Only small dimensional changes occur during processing since the matrix fills the pores within the preform without disturbing the reinforcement. A disadvantage is the presence of residual aluminum phase (~ 5-10%) in the matrix that must be removed if the part is to be used above the melting point of aluminum.

In LSI, liquid silicon, or one of its lower melting point alloys, is used to infiltrate a fiberous preform to form a silicon carbide matrix. Before infiltration, a fine grained silicon carbide particulate is slurry cast into the fiber preform. After removal of the slurry carrier liquid, melt infiltration is usually done at 2550° F, or higher, and is usually complete within a few hours. The liquid silicon bonds the silicon carbide particulates together and forms a matrix that is somewhat stronger and denser than that obtained by CVI. Since the resulting matrix can contain up to 50% unreacted silicon, the long-term use temperature is limited to about 2200° F. The liquid metal infiltration processes have several advantages: (1) they produce a fairly dense SiC based matrix with a minimum of porosity; (2) the processing time is shorter than for most ceramic matrix composite fabrication processes; and (3) the dense and closed porosity on the surface can often eliminate the need for a final oxidation resistant coating. The major disadvantage is the high temperatures required for liquid silicon infiltration that exposes the fibers to possible degradation, due to the high temperatures and the corrosive nature of liquid silicon.

Recommended Reading

[1] Handbook on Continuous Fiber-Reinforced Ceramic Matrix Composites, Ceramics Information Analysis Center, 1995.

[2] Chawla, K.K., Ceramic Matrix Composites, Chapman & Hall, 1993.

References

[1] Buckley, J.D., "Carbon-Carbon Composites", in Handbook of Composites, Chapman & Hall, 1998, pp. 333-351.

[2] Amateau, M.F., "Ceramic Composites", in Handbook of Composites, Chapman & Hall, 1998, pp. 307-332.

[3] DiCarlo, J.A., Dutta, S., "Continuous Ceramic Fibers for Ceramic Matrix Composites", in Handbook on Continuous Fiber-Reinforced Ceramic Matrix Composites, Ceramics Information Analysis Center, 1995, pp. 137-183.

[4] 3M Nextel Ceramic Textiles Technical Notebook, 3M Ceramic Textiles and Composites, 2003.

[5] Marzullo, A., "Boron, High Silica, Quartz and Ceramic Fibers", in Handbook of Composites, Chapman & Hall, 1998, pp. 156-168.

[6] "State of the Art in Ceramic Fiber Performance", in Ceramic Fibers and Coatings: Advanced Materials for the Twenty-First Century, The National Academy of Sciences, 1998, pp. 20-36.

[7] DiCarlo, J.A., Yun, H.M., "New High-Performance SiC Fiber Developed for Ceramic Composites", NASA-Glenn Research & Technology Reports, 2002.

[8] Luthra, K.L., Corman, G.S., "Melt Infiltrated (MI) SiC/SiC Composites for Gas Turbine Applications", GE Research & Development Center, Technical Information Series, 2001.

[9] Naslain, R.R., "Ceramic Matrix Composites: Matrices and Processing", in Encyclopedia of Materials: Science and Technology, Elsevier Science Ltd, 2000.

[10] Zolandz, R., Lehmann, R.L., "Crystalline Matrix Materials for Use in Continuous Filament Fiber Composites", in Handbook on Continuous Fiber-Reinforced Ceramic Matrix Composites, Ceramics Information Analysis Center, 1995, pp. 111-136.

[11] DiCarlo, J.A., Bansal, N.P. "Fabrication Routes for Continuous Fiber-Reinforced Ceramic Composites (CFCC)", NASA/TM-1998-208819, 1998.

[12] Lange, F.F., Lam, D.C.C., Sudre, O., Flinn, B.D., Folsom, C., Velamakanni, B.V., Zok, F.W., Evans, A.G., "Powder Processing of Ceramic Matrix Composites", Material Science and Engineering, A144, 1991, pp. 143-152.

[13] Chawla, K.K., "Processing of Ceramic Matrix Composites", in Ceramic Matrix Composites, Chapman & Hall, 1993, pp. 126-161.

[14] Becker, P.F., Tiegs, T.N., Angelini, P., "Whisker Reinforced Ceramic Composites", in Fiber Reinforced Ceramic Composites: Materials, Processing and Technology, Noyes Publications, 1990, pp. 311-327.

[15] Lewis III, D., "Continuous Fiber-Reinforced Ceramic Matrix Composites: A Historical Overview", in Handbook on Continuous Fiber-Reinforced Ceramic Matrix Composites, Ceramics Information Analysis Center, 1995, pp. 1-31.

[16] French, J.E., "Ceramic Matrix Composite Fabrication and Processing: Polymer Pyroly-sis", in Handbook on Continuous Fiber-Reinforced Ceramic Matrix Composites, Ceramics Information Analysis Center, 1995, pp. 269-299.

[17] Cullum, G.H., "Ceramic Matrix Composite Fabrication and Processing: Sol-Gel Infiltration", in Handbook on Continuous Fiber-Reinforced Ceramic Matrix Composites, Ceramics Information Analysis Center, 1995, pp. 185-204.

[18] Mah, T., Yu, Y.F., Hermes, E.E., Mazdiyasni, K.S., "Ceramic Fiber Reinforced Metal-Organic Precursor Matrix Composites", in Fiber Reinforced Ceramic Composites: Materials, Processing and Technology, Noyes Publications, 1990, pp. 278-310.

[19] Naslain, R., "Design, Preparation and Properties of Non-Oxide CMCs for Application in Engines and Nuclear Reactors: An Overview", Composite Science and Technology, Vol. 64, 2004, pp. 155-170.

[20] Lowden, R.A., Stinton, D.P., Besmann, T.M., "Ceramic Matrix Composite Fabrication and Processing: Chemical Vapor Infiltration", in Handbook on Continuous Fiber-Reinforced Ceramic Matrix Composites, Ceramics Information Analysis Center, 1995, pp. 205-268.

[21] Lackey, W.J., Starr, T.L., "Fabrication of Fiber-Reinforced Ceramic Composites by Chemical Vapor Infiltration: Processing, Structure and Properties", in Fiber Reinforced Ceramic Composites: Materials, Processing and Technology, Noyes Publications, 1990, pp. 397450.

[22] Fareed, A.S., "Ceramic Matrix Composite Fabrication and Processing: Directed Metal Oxidation", in Handbook on Continuous Fiber-Reinforced Ceramic Matrix Composites, Ceramics Information Analysis Center, 1995, pp. 301-324.

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Chapter 11

Structural Assembly

Assembly represents a significant portion of the total manufacturing cost. Assembly operations are labor intensive and involve many steps and, as shown in Fig. 11.1, can represent as much as 50% of the total delivered part cost.1 For example, the wing shown in Fig. 11.2 requires:

(1) A framing operation in which all of the spars and ribs must be located in their proper location and connected together with shear ties.

(2) Each skin must then be located on the substructure, shimmed, holes drilled, and fasteners installed. During and after skin installation there are various sealing operations that must be performed.

(3) The final wing torque box must have the leading edges, wing tips, and control surfaces assembled.

This brief description is a gross over-simplification of the complexity involved in assembling a large structural component. The number of mechanical fasteners in a typical fighter aircraft might be in the range of 200 000-300 000, while a commercial airliner or transport aircraft can have as many as 1 500 0003 000 000 fasteners, depending on aircraft size. A hole has to be drilled for each of these fasteners and then the fastener has to be installed in each hole. In this chapter, the basic assembly operations will be covered with an emphasis on hole preparation and the types of mechanical fasteners used in aircraft structures.

11.1 Framing

Framing operations, in which the substructure is located and fastened in its proper location, has made significant progress since the mid-1960s. In the 1960s and 1970s, substructure was primarily manually located using some hard tool located positions, usually supplemented with large pieces of clear plastic film (Mylar's) scribed with hole pattern locations. The design, tooling, and fabrication databases were not necessarily coordinated with each other, which created a lot

Materials Fasteners and and

Materials Fasteners and and

Fig. 11.1. The High Cost of Assembly Operations1
Fig. 11.2. Assembly Complexity Source: The Boeing company
Fig. 11.3. Laser Projection Location Source: Laser Projection Technologies, Inc.

of variability and poorly fitting parts. During the 1980s and 1990s, there was less use of Mylar's and a greater reliance on hard tooling to position parts. However, the extensive use of hard tooling to locate parts increased the nonrecurring investment required at the start of a new program. With the advent of solid modeling and electronic master models in the 1990s, a process called determinant assembly emerged. In determinant assembly, coordinated undersize fastener holes are drilled in the parts during fabrication. These holes are then used to position the parts during assembly, eliminating the need for hard tooling locators. Another recent development is the use of laser projection units for establishing part and hole location. A typical application for a laser projection system is shown in Fig. 11.3.

11.2 Shimming

Prior to starting hole drilling and fastener installation, it is important to check all joints for the presence of gaps. The presence of gaps can unnecessarily preload metallic members when fasteners are installed, a condition that can initiate premature fatigue cracking and even stress corrosion cracking of aluminum. However, gaps in structure containing composites can cause even more serious problems than in metallic structure. Since composites do not yield, and are more brittle and less forgiving than metals, excessive gaps can result in delaminations when they are pulled out during fastener installation. The composite is put in bending due to the force exerted by the fastener drawing the parts together and can develop matrix cracks and/or delaminations around the holes. Cracks and delaminations usually occur on multiple layers through the thickness and can adversely affect the joint strength.2 Gaps can also trap metal chips and contribute to backside hole splintering. If the skin is composite and the substructure is metal, and if an appreciable gap is present during fastener installation, the composite skin will often crack and delaminate. If both the skin and substructure are composite, then cracks can develop in either the skin or substructure, or both. Substructure cracking often occurs at the radius between the top of the stiffener and the web.

To prevent unnecessary preloading of metallic structure, and the possibility of cracking and delaminations in composite structure, it is important to measure all gaps and then shim any gaps greater than 0.005 in. Liquid shim, which is a filled thixotropic adhesive, can be used to shim gaps between 0.005 and 0.030 in. If the gap exceeds 0.030in., then a solid shim is normally used, but engineering approval is often required for a gap this large. Solid shims can be made from solid metal, laminated metal that can be peeled to the correct thickness, or composite. When selecting a solid shim material, it is important to make sure there is no potential for galvanic corrosion within the joint.3

Liquid shimming can be accomplished by first drilling a series of undersize holes in the two mating surfaces, for installing temporary fasteners, to provide a light clamp-up during the shimming process. The liquid shim is usually bonded to one of the two mating surfaces. The surface that will be bonded should be clean and dry, to provide adequate adhesion. Composite surfaces should be scuff sanded. The other surface is covered with release tape or film. After the liquid shim is mixed, it is buttered onto one surface and the other surface is located, and then clamped up with mold-released temporary fasteners. The excess or squeeze-out is removed prior to gellation, which usually occurs within an hour of mixing. After the shim material is cured, typically for about 16 h, the part is disassembled and any voids or holes in the shim are repaired. After the repair, the parts are assembled.

11.3 Hole Drilling

It should be noted that there are some differences between fighter aircraft and commercial passenger aircraft. Fighter aircraft designs are highly tailored to performance and loads, leading to a lot of thickness variations in the skins and substructure to reduce weight. This results in a wide variety, but limited numbers, of fastener types, grip lengths, and diameters. Due to the smaller size of a fighter airframe, there are more areas of limited access during assembly. On the other hand, larger commercial aircraft have much more fastener commonality with regards to type, grip length, and diameter, but also, due to their size, many more fasteners. Skins and substructure tend to be more uniform in thickness. Limited access is not as much of a problem, but the shear size of the parts makes them difficult to handle. There are many types of drill motors and units that can be used to drill structures, but they can be broadly classified as either hand, power feed, automated drilling units, or automated riveting equipment.

11.3.1 Manual Drilling

Manual, or free hand, drilling using hand-held drill motors, such as the one shown in Fig. 11.4, has the least chance of making a close tolerance hole (+0.003/ — 0.000in.). The only real control is the drill speed (rpm). It is up to the operator to make sure the drill: (1) is located in the proper location; (2) is perpendicular to the surface; and (3) is fed with enough pressure to generate the hole, but not too much pressure to damage the hole. Although free hand drilling is obviously not the best method, it is frequently used, because it requires no investment in tooling (i.e., drill templates), and, in many applications, where access is limited, it may be the only viable method. A typical tight access situation is shown in Fig. 11.5, where a mechanic is installing collars on Hi-Lok fasteners. For tight access areas, right-angle drill motors are available. If free hand drilling is used, it is recommended that the operators use a drill bushing, or tri-pod support, to insure normality, and that they be provided with detailed written instructions for hole generation and inspection.

Manual hole drilling during assembly is often done by drilling undersize holes (pilot holes), installing temporary fasteners to the hold the parts together, and then bringing the holes up to full size after all of the pilot holes are drilled. Pilot holes are usually drilled with small diameter drill bits (e.g., 0.90-0.125 in.). Hole diameters for aerospace structures nominally range from 0.164 to 0.375 in., with the predominate hole sizes being 0.188 and 0.250 in. diameter.4 During drilling, it is important to use a sharp drill bit: (1) a sharp drill bit will not wander as

Fig. 11.4. Typical Free-hand Drill Motor Source: Cooper Power Tools
Fig. 11.5. Limited Access Fastener Installation Source: The Boeing Company

easily as a dull one, (2) drilling is faster with a sharp bit, and (3) lower forces are required, minimizing the possibility of injury or part damage. When drilling multi-material stack-ups, it is important to make sure they are securely clamped together. A typical assembly with temporary fasteners installed to hold the parts in position and provide clamp-up is shown in Fig. 11.6.

The speeds used in manual drilling depend on both the materials and their thickness. Typical maximum speeds for aluminum range from 1000 rpm for 0.50 in. holes up to 5000 rpm for 0.16 in. holes. For titanium, they range from 150 rpm for 0.50 in. holes up to 700 rpm for 0.16 in. holes. As the material gets thicker, the maximum speed goes down, and as the material gets harder, the maximum speed allowed goes down. When drilling through stack-ups of dissimilar metals (e.g., aluminum and titanium), it is necessary to use a drill motor with the speed adjusted for the harder material.

After all the holes have been drilled, the assembly should be taken apart, and the holes deburred on both surfaces. Holes should only be deburred with

Fig. 11.6. Panel with Temporary Fasteners Source: The Boeing Company

a deburring tool, or a drill bit of a larger size. Holes should be deburred by hand due to the danger of hole enlargement if a power drill is used. It is also important not to go too deep as shown in Fig. 11.7, as a knife edge can result. During all hole drilling operations, it is important that proper edge distances are maintained, to insure that the skins do not fail due to inadequate shear strength. The engineering drawing should specify edge distances but they are normally around 2-3D, where D is the hole diameter.

Deburred Holes

Fig. 11.7. Deburring of Drilled Holes

Knife Edge

Section Through Sheet

Fig. 11.7. Deburring of Drilled Holes

Hole drilling of composites is more difficult than in metals, again due to their relatively low sensitivity to heat damage, and their weakness in the through-the-thickness direction. Composites are very susceptible to surface splintering (Fig. 11.8), particularly if unidirectional material is present on the surface. Note that splintering can occur at both the drill entrance and exit side of the hole.6 As shown in Fig. 11.9, when the drill enters the top surface, it creates peeling forces on the matrix as it grabs the top plies. When it exits the hole, it induces punching forces that again creates peel forces on the bottom surface plies. If top surface splintering is encountered, it is usually a sign that the feed

Drill Induces Peeling Forces Drill Induces Punching Forces on Top Plies During Entry on Bottom Plies During Exit

Fig. 11.9. Drilling Forces on Composite Laminate2

Drill Induces Peeling Forces Drill Induces Punching Forces on Top Plies During Entry on Bottom Plies During Exit

Fig. 11.9. Drilling Forces on Composite Laminate2

rate is too fast, while exit surface splintering indicates that the feed force is too high.3 It is common practice to cure a layer of fabric on both surfaces of composite parts, which will largely eliminate the hole splintering problem, i.e. woven cloth is much less susceptible to splintering than unidirectional material. When drilling composites, a back-up material such as aluminum or composite, clamped to the backside, will frequently help to prevent backside hole splintering. Coolant is normally not used for carbon/epoxy laminates that are 0.250 in. thick or thinner. When drilling composites dry, operators should be provided with vacuum capability to suck up the dust and should always wear eye protection and a respirator.

Since epoxy matrix composites will start to degrade if heated above 400° F, it is important that heat generation be minimized during drilling. Typical drilling parameters are 2000-3000 rpm at feed rates of 0.002-0.004 in. per revolution (ipr), although this will vary depending on the drill geometry and the type of equipment used. Thermocouples and heat sensitive paints are often used during drilling parameter development tests to monitor the heat generated. Drilling parameters for composite-to-metal stack-ups are often controlled more by the metal than the composite. For example, when drilling C/E-to-aluminum, a speed of 2000-3000 rpm with a feed rate of 0.001-0.002 ipr might be used, while a stack-up of C/E-to-titanium would require a slower speed (e.g., 300-400rpm) and a higher feed rate (0.004-0.005 ipr). Titanium (Ti-6Al-4V) is also very sensitive to heat build-up (hence the lower speed) and tends to rapidly work harden if light cuts are used (hence the higher feed rate).

To help reduce the variability in manual hole drilling, some manufacturers have produced detailed written instructions covering specific hole drilling operations, and provide kits for the mechanics, which have all of the correct tools needed for a specific operation.

11.3.2 Power Feed Drilling

Power feed drilling is much preferred to hand drilling. In power feed drilling, the drill unit is locked into a drill template that establishes both hole location and maintains drill normality. In addition, once the drilling operation starts, the unit is programmed to drill at a given speed and feed. Some units, such as the one shown in Fig. 11.10, can be programmed for different peck cycles. All of these controls lead to much better and more consistent hole quality, particularly when drilling composite-to-metal stack-ups. A typical peck drilling cycle for a 3/16 in. diameter hole through C/E-to-titanium would be a speed of 550 rpm with a feed rate of 0.002-0.004 ipr and 30-60 pecks per inch of thickness.7

When drilling into composite-to-metal stack-ups, a phenomenon called back counterboring can occur. As shown in Fig. 11.11, as the metal chips (i.e., aluminum or titanium) travel up the flutes, they tend to erode the softer liquid shim and composite matrix material, causing eroded and oversize holes. Back

Fig. 11.10. Power Feed Peck Drill Source: Cooper Power Tools

counterboring can be minimized by (1) eliminating all gaps, (2) using a drill geometry that produces small chips, (3) changing speeds and feeds, (4) providing better clamp-up, (5) reaming the hole to final diameter after drilling, or (6) by peck drilling.3

Peck drilling (Fig. 11.12) is a process in which the drill bit is periodically withdrawn to clear the chips from the flutes. Peck drilling is used almost exclusively when drilling composite-to-titanium stack-ups, due to the back counterboring potential of the hard titanium chips. The process also greatly reduces the heat build-up that can rapidly occur when drilling titanium.

11.3.3 Automated Drilling

For high volume hole generation, automated drilling equipment can be designed and built for specific applications.8 Being large and sophisticated machine tools,

Fig. 11.12. Peck Drilling3
Fig. 11.13. Automated Wing Drilling System Source: The Boeing Company

these units are expensive, so the number of holes drilled and the number of units produced needs to be large enough to justify the equipment investment. Examples of two of these large units are shown in Figures 11-13 and 11-14. These machines are extremely rigid and allow for accurate hole location and normality. They are NC so there is no need for drill templates. The one shown has a vision system that can scan the substructure, and software that will then adjust the hole location to match where the substructure is actually located, versus design nominal. All drilling parameters are automatically controlled with the capability to change speeds and feeds, when drilling through different materials. Due to the thick stack-ups that must be drilled in a wing, a water soluble flood or mist coolant is usually used during the hole drilling operations. All drilling data is automatically recorded and stored for quality control purposes. The drill holders contain bar codes that must match the

Four Independent Drill Columns

Drill Column

Fig. 11.14. Automated Wing Drilling System Source: PaR Systems, Inc.

Drill Column

Fig. 11.14. Automated Wing Drilling System Source: PaR Systems, Inc.

drilling program to make sure the correct drills are used for the correct holes. These machines can also install temporary fasteners to clamp the skins to the substructure during drilling, and frequently use integral drill-countersink cutters that drill the hole and then continue to countersink it, during the same operation.

The current trend in industry is to replace these large installations with smaller more flexible units.910 An off-the-shelf commercial robot, with some modifications and a special drilling end effector, is used to drilling holes in the control surface shown in Fig. 11.15. Another approach is to integrate drilling

Fig. 11.15. Robotic Drilling of Control Surface Source: The Boeing Company

units into the assembly fixture, an approach called Numerically Controlled Drill Jigs (NCDJ). Examples of these types of relatively low cost units are shown in Fig. 11.16 for a forward fuselage and Fig. 11.17 for a fighter aircraft outer wing.

11.3.4 Automated Riveting Equipment

A typical piece of automated riveting equipment is shown in Fig. 11.18. These machines will drill the hole, inspect the hole, select the correct grip length of rivet, install sealant on the rivet (if required), and then install the rivet by squeezing. This equipment is available in a wide variety of sizes, ranging from small units to very large computer numerically controlled (CNC) units, capable of installing stringers on full size commercial aircraft wing skins. Being an automated process, the quality of drilling and fastener installation is better and more consistent than with hand methods. There are also units that will install pin and collar fasteners, such as Lockbolts, where the collar is automatically swaged onto the collar portion of the pin. An example of one of these larger automated drill and fastening systems is shown in Fig. 11.19.

Forward Fuselage Tooling

Drilling Operation

Fig. 11.16. Forward Fuselage NCDJ Drilling System9

Drilling Operation

Fig. 11.16. Forward Fuselage NCDJ Drilling System9

11.3.5 Drill Bit Geometries

Many variations of twist drills (Fig. 11.20) are used in drilling metallic structure. Since specific drill bit geometries can influence both hole quality and the quantity of holes drilled, many geometries are proprietary to the various

aerospace manufacturers. Examples of typical variations in the standard twist drill include step drills that drill an undersize hole and then a final hole size in one pass, and drill-reamers in which ahole is drilled and then reamed in the same pass. When drilling aluminum, standard high speed steels, such as M2 or M7, give satisfactory drill life. For harder materials, such as titanium, the cobalt grades of high speed

Fig. 11.18. Medium Size Automatic Riveter Source: The Boeing Company
Fig. 11.19. Large Automated Fastening System Source: The Boeing Company

Point Cutting

Angle Edge

Chisel Edge

Point Cutting

Angle Edge

Chisel Edge

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