1994 toyota 4runner service repair manual software
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1994 toyota 4runner service repair manual softwareThen a review of applications is presented, starting with early examples from the 1930s and the Second World War. The growth in the use of sandwich materials in civil and military applications is then developed. Recent research and innovations conclude the paper. Published by Elsevier B.V. Recommended articles No articles found. Citing articles Article Metrics View article metrics About ScienceDirect Remote access Shopping cart Advertise Contact and support Terms and conditions Privacy policy We use cookies to help provide and enhance our service and tailor content and ads. By continuing you agree to the use of cookies. The faces are usually made of laminated polymeric based composite materials, and typically, the core can be a honeycomb type material, a polymeric foam or balsa wood. The faces and the core are joined by adhesive bonding, which ensures the load transfer between the sandwich constituent parts. The result is a special laminate with very high bending stiffness and strength to weight ratios. These limits are influenced by many material, structural, damage, and loading variables. Large ADLs are desirable to minimize the maintenance burden of the aircraft operator. The methods can be used to explicitly predict the residual strength of a known damage, or to determine Allowable Damage Limits (ADLs) for inclusion in the Structural Repair Manual (SRM). The approach uses damage metrics compatible with airline operator inspection methods, and addresses all in-service damage types for a wide range of configurations typically found in fairing and flight-control-surface applications on current Boeing airplanes. However, some generalized strength results allow estimating the residual strength response of other materials, as might be desired in preliminary design. The methodology represents an improvement over past methods through higher-fidelity analyses and consideration of damage severity.http://www.ylasavonrasti.net/tiedostot/download-kiv-7m-manual.xml
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This latter consideration avoids a major conservatism often included in past approaches, where delaminations, dents, and punctures are all treated as holes. The semi-empirical prediction method is based on a power-law relationship between strength and damage size, and relies heavily on a large database of notched and impact damaged strengths of thin-gage honeycomb configurations collected at Boeing since the 1980’s. Continuous functions were developed to capture the effects of layup and damage severity on residual strength (using orthotropic stress concentration factors and damage-diameter-to-depth ratios, respectively), leveraging data trends from all materials. Regression analyses were then used to characterize the response of each material.Statistics, environment, core thickness, and finite-width effects are addressed through the use of correction factors. Additional approaches are employed to address facesheet thickness, combined loading, non-circular damage, hybrid layups, facesheet springback, and multi-site damage. These correction factors and approaches leverage data, where available, and use conservative assumptions where uncertainties exist. Existing analysis methods form the basis for predicting (a) sublaminate buckling, (b) facesheet fracture due to the buckled sublaminate, and (c) static delamination growth, all as a function of damage size and delamination depth. These predictions are compared to the applied loading to preclude static failure and fatigue delamination growth. Conservative techniques are employed to address limitations of in-service inspection methods relative to identifying delamination depth and the existence of multiple delaminations. A number of damage size and spacing measurements are necessary to support the above methods. SRM allowable damage data for one zone of a specific aircraft part include one plot addressing single-site damage, while a second plot provides a correction factor for multi-site damage scenarios.http://goldenlionjalisco.com/archivos/download-knex-manuals.xml The left-hand portion of the curve is based on the impact and hole residual strength predictions. The curve is generated by determining the damage size that results in failure at the applied Ultimate strain from the residual strength curves for the specific layup. The right-hand portion of the curve reflects the results for delaminations with no dent. These constant values are a function only of delamination depth. Each line represent the smallest of the damage sizes associated with the failure modes related to sublaminate buckling (facesheet fracture, static and fatigue delamination growth) for the applied strains and the delamination depth indicated. Evaluations of impact damage surveys indicate that springback occurs only in less-severe damage scenarios, and that the severity level below which springback might occur is a function of facesheet thickness. The residual strength at this severity threshold is used as an upper limit for less severe damages. However, damage depths associated with facesheet springback might also relax over time when subjected to environmental and mechanical cycling associated with service. Damages where this is considered to be a possibility are therefore conservatively treated as a delamination with no dent, except in cases where intermediate inspections can and will be performed to monitor the damage for growth. Specifically, isotropic stress concentration factors were found to reasonably predict the interactions observed in test data, and are therefore used to define strength knock-down factors based on damage spacing, relative damage size, and loading orientation. The correction factors for damage size are determined from these strength-based knock-downs, and are dependent on layup due to the dependence of the notch sensitivity on layup. Severity differences between damages within a pair are treated conservatively by assuming both damages are of the greater severity of the two damages. The operator obtains the required damage size and spacing measurements. If the maximum size of the actual damage is less than the ADL, and all spacing requirements are met, it is acceptable. Otherwise, the damage must be repaired. The severity of both damages in a pair are calculated, and the smaller of the two is used to determined a single-site ADL from the upper curve. The multi-site ADLs are determined for each damage as the product of the single-site ADL and the respective MDCF. If the maximum size of the actual damages are both less than their respective ADLs, and all spacing requirements are met, they are acceptable. Otherwise, both damages must be repaired. If you are unable to make a purchase, please try again after 7:00AM EDT. Thank you for your patience. Search for more papers by this author Published Online: 8 Oct 2018 Sections PDF PDF Plus Abstract Over the past few years, many studies have been focused on finding highly integrated composite structures that can replace such sandwich designs. The motivation is given mainly because of commercial aspects, which from a structural point of view are difficult to justify. Sandwich structures have the intrinsic property of high bending stiffness and are therefore perfectly suitable for the plate-like design of an aircraft spoiler. However, monolithic structures can be produced with a lower cost, and their subcomponents can be easily integrated in a single part. Two monolithic alternatives, which meet all structural requirements, are considered with respect to applied materials, manufacturing processes, and manufacturing costs. The established designs are compared with a standard sandwich spoiler, and thus a holistic view is provided on the selection of designs, processes, and materials for future aircraft spoilers. I. Introduction S poilers of large commercial aircraft are often realized in composite sandwich design. The spoiler body is made of carbon-epoxy plies. Those cover an aramid honeycomb core.http://daniela-vashiron.com/images/conrep-hp-manual.pdf Nowadays, research is focused on monolithic highly integrated composite structures that substitute such sandwich designs, mainly because of commercial aspects. From a structural point of view, this is difficult to justify. However, monolithic structures can be produced with a lower cost and their subcomponents can be easily integrated in a single part. In the literature, only few studies are focused on monolithic spoilers. For one of those a demonstrator was finally manufactured. In both studies, the previously mentioned monolithic spoiler is used as a best practice example. If the distance between two radial ribs becomes too large, a new branch is introduced. However, the development focuses on metal sintering via rapid prototyping technology. At the same time they have approximately the equal mass compared with a sandwich spoiler. However, in the development of structures made from fiber-reinforced plastic, the areas of design, manufacturing processes, and materials are closely connected. A change on one area can have a major impact on the other areas. Therefore, the developed concepts are transferred to two monolithic design alternatives of an existing sandwich spoiler. The requirements for the monolithic alternatives are same aerodynamic surface and interface geometry, same sizing criteria, similar global stiffness behavior, and similar mass compared with the sandwich spoiler. Instead of the materials used for the sandwich spoiler, newer aircraft qualified materials and their material properties are applied. The design space is given by the volume of the sandwich spoiler. In this publication, both design alternatives are considered in terms of applied materials, manufacturing processes, and the resulting manufacturing costs. Compared with the standard sandwich spoiler, they provide a holistic view on alternative designs of future aircraft spoilers. This paper is divided into three sections. First, the sandwich design and its monolithic alternatives are explained, a review about the different approaches toward the monolithic designs is given, and the critical design criteria for each spoiler type are highlighted. Second, manufacturing processes and applied materials for the individual designs are sketched. Design details heavily depend on selected processes and materials. Thus some of them are presented in this section too. Third, a comparison of manufacturing costs between the sandwich spoiler and the monolithic spoilers is presented. Changing the design from sandwich to monolithic does not automatically result in less manufacturing costs. Typically, complexities are shifting, for example, from bond assembly to tooling. Therefore, a trade-off between all aspects of the possible design options is needed to finally find the most economical solution. However, before details are provided, the functionality of aircraft spoilers, the production of sandwich structures, and the production of monolithic structures are summarized. The design alternatives presented do not claim to be the only solutions; for example, less-expensive manufacturing processes can be applied also on sandwich spoilers, if foam cores are used instead of the state-of-the-art honeycomb cores. However, foam core solutions are not considered, because the main drawbacks of foams are similar to honeycomb solutions. Spoilers are deployed for several tasks. Deflected during flight, they act like an air brake and reduce the aircraft speed (speed brake). Immediately after landing the spoilers are extended to their maximum deflection (ground spoiler). Again the drag is increased; however, the main effect is that the wing surfaces located behind the spoilers (e.g., the flaps) are prevented to provide lift. Therefore, the contact forces between the wheels and the runway increase, which increases the braking efficiency. When spoilers are applied as roll spoiler to support spiraling, they are only deflected on the side where the aileron is shifted upward. The one-sided lift reduction causes the aircraft to roll. At the same time, due to the one-sided increased drag, the aircraft starts to yaw to the side where the spoilers are deflected. In an aircraft with electronic flight control, spoilers can be used to reduce the loads acting on the wing. This is known as gust load alleviation. Vertical wind gusts increase the acting air-load. By fast deflections upon the occurrence of wind gusts the air-load can be alleviated, which allow a lighter wing design. Recently developed aircraft droop the spoilers slightly when the flaps drive to their extended position (e.g., the Airbus A350 and the Boeing 787). The reason is to keep a defined gap between the trailing edge of the spoiler and the extended flap to allow a controlled airflow from the lower wing side to the upper wing side. Here, the drooped spoilers function as a high lift device. Fig. 1 Schematic layout of an aircraft wing. B. Manufacturing Processes Sandwich structures such as spoilers are usually built with resin pre-impregnated carbon fiber material and an aramid honeycomb core. Because of their complexity, the assembly of the individual ply and the honeycomb core segments is carried out in several curing steps. Curing is performed in an autoclave, which is costly, because high pressures and temperatures are applied in a nitrogen-filled environment. Monolithic designs allow a variety of semifinished fiber products and manufacturing processes, that is, resin transfer molding (RTM) and resin infusion (RI). For both, preforms made from dry carbon fibers are saturated with epoxy resin during the curing process. For RI, vacuum is used to pull the resin through the preform. In case of RTM the resin is injected under high pressures. The curing process can be performed in a simple oven or with heatable molds. The use of preforms has a further advantage. Subcomponents can be integrated to one complex monolithic structure, which can be manufactured in a single layup and curing step. The spoilers of those aircraft are developed and produced by FACC Operations GmbH, Ried im Innkreis, Austria (FACC). The overall dimensions of the spoiler in millimeter are 2400 ? 800 ? 150. The magnitude of the acting loading is given by the forces acting on the actuator rod, which is attached to the actuator lugs and is used to extend or retract the spoiler. The spoiler is designed such that it can withstand in extended position aerodynamic loads causing a compression force of 180 kN in the actuator rod. Tension loading of the actuator occurs when the spoiler is forced against the flap to prevent lift off due to aerodynamic loads. In this case an actuator malfunction is considered additionally, and therefore the spoiler has to withstand a tension force of 130 kN in the actuator. The design space, the interface geometry, the sizing criteria, the structural behavior, and the structural mass form the basis for the monolithic design alternatives. In the following the sandwich design as reference and the two derived monolithic spoiler designs are presented. A. Sandwich Spoiler: The Reference Design In Fig. 2 the schematic design of the sandwich spoiler is shown. The spoiler is connected via hinges and actuator bearings to the wing structure. The spoiler body is produced by a Nomex honeycomb core, which is covered by the upper skin laminate and the lower skin laminate. At the leading edge the sandwich is closed with a C-spar. At the trailing edge a laminated wedge is applied. All laminates are produced from epoxy resin pre-impregnated carbon plies. In spreadwise direction the sandwich structure is simply closed by a ramp at the inboard edge and the outboard edge. The interface features are realized with one bracket in the center of the leading edge and two smaller brackets at booth outer corners. The center hinge (CH) bracket, made from carbon-fiber-reinforced plastic (CFRP), is produced in an RTM process. Subsequently, after few process steps, the center hinge is integrated in the bond assembly of the sandwich structure. The CH consists of four lugs, two outer ones used for the main hinge bearings and two center lugs, where the actuator is attached. The global stiffness of the structure in the present case is a result of above requirements and not a requirement itself. Rather large displacements are acceptable as long as the structural strength is satisfied. The structural strength is calculated through application of the maximum stress criteria for all ultimate load cases. If unidirectional (UD) carbon fiber layers are used, the Tsai-Wu first ply failure criteria is applied for matrix failure for limit load cases. Stability is checked by linear buckling analysis. Here, the reserve factor against buckling has to be larger than one, for limit load. In respect to the design philosophy a combination of the different approaches is applied. Fail-safe design is applied for the hinge brackets at the inboard and the outboard edge (Fig. 2 ). If one of these brackets fails, the structure must be still able to carry limit load. Before bracket failure, the load is transferred directly into the wing structure. After bracket failure, the load is carried via the center hinge into the wing structure. The composite laminates are designed with the damage tolerance approach. If in-plane strains do not exceed a certain value, for example, 0.0045 for longitudinal and transverse tension, endurance life is given. This is valid even when barely visible damages are present. The most influencing criteria for the sandwich structure are the maximum strain criteria for damage tolerance. With the exception of special cases it can be assumed that structural strength is satisfied if the damage tolerance criteria in terms of a maximum strain allowable is fulfilled. Sandwich structures are naturally uncritical for global buckling. Local buckling in the cell walls is not a significant criterion. B. Monolithic Spoiler I: System of Integrated Stiffeners The idea for the first monolithic design alternative is to replace the sandwich as load carrying structure by an air loads carrying skin and a support structure, which is based on a system of stiffeners. However, the application of sandwich design is perfectly suitable for bending loaded structures such as spoilers. Thus, omitting the lower skin has a major impact on structural stiffness and strength, which must be alleviated to establish a design of the approximately equal mass. Based on the given design space an objective function is formulated together with constraints, which reflects the design criteria of the A330 and A340 aircraft spoiler as described in the previous section. It forces the solver toward solutions showing a distribution of ribs. The topology optimization results are interpreted to the design presented in Fig. 3. Here, the spoiler is produced by a simple upper skin and a rather complex backbone, which are bonded together. The backbone structure is designed as a system of integrated omega stiffeners, which are distributed over the skin area to support the pressure-loaded skin. The CH bracket is integrated into the backbone structure, which is produced in one ply layup step. Again, aluminum brackets are installed next to the inner and outer edge of the spoiler. Those are slightly adapted to meet the new geometric features of the spoiler body. Fig. 3 Schematic design of monolithic spoiler I based on an integrated stiffener system. Compared with the sandwich structure, two design criteria become critical. The maximum strain criteria still dominate the design. Additionally, the reserve against buckling is critical, which is reasonable due to the unsupported skin sections between the omega stiffeners. C. Monolithic Spoiler II: A Closed Box Design For the second monolithic design concept the original idea of a sandwich design is substituted by chordwise-oriented ribs. Low density is one of the motivations for using sandwich material. It is used to produce a distance between two stiff skins. Further, the top and the bottom skins carry the tensional and compressive stresses and the sandwich core carries the shear stresses. In this way, panels with high bending stiffness can be produced. This advantage is used for the design of monolithic spoiler II. Thus, the basic components of the sandwich spoiler (Fig. 2 ) are adopted, and just the sandwich core is replaced by simple shells as presented in Fig. 4. The four lugs of the CH are built by four of those shells, and thus a well-suited load introduction is realized. To transfer shear forces the structure is closed with C-spar sections. Additionally, these C-spar sections prevent free edge buckling. The number of shells and the distance between them are heavily influenced by the buckling criteria, local displacements due to the air-loads-carrying upper skin, and the laminate thickness. Fig. 4 Schematic design of monolithic spoiler II based on chordwise-oriented ribs. D. Comparison of Weight The weight of monolithic spoiler I is about 8 higher and the weight of monolithic spoiler II is about 14 less compared with the sandwich spoiler. However, these values are derived from different development levels. The weight of monolithic spoiler I is derived from a design, which is at a high development level, e.g. demonstrator parts were produced, static load tests were conducted and a second weight optimization loop was already performed. The design of monolithic spoiler II is still in the concept phase, and some increase in weight may happen because of detail design. III. Production Processes and Design Details In this section, the manufacturing processes of the individual designs are described. Thus some of them are presented in this section too. A. Sandwich Spoiler The first step in manufacturing the sandwich spoiler presented in Fig. 2 is to produce the center hinge bracket. Because of its complex geometry and its tight tolerances, this is done applying an RTM process. Therefore, dry carbon fiber weaves are assembled to a preform. The preform is then placed in the RTM press where all surfaces are covered with heatable mold blocks. The resin is injected and the CH is cured. For the produced composite part a CNC-machining step is required, nondestructive testing (NDT) is needed, and then the bonding surfaces must be prepared for further processing. In the second step, the layup of the C-spar takes place. Therefore, the center hinge is placed in the layup mold for cobonding. Then, the prepreg layers are applied to form the C-spar laminate, which is cured in an autoclave. First at the final ply layup step, which is visualized in Fig. 5, the upper skin prepreg plies are placed onto the layup mold. Those are topped with film adhesive. Then the C-spar, which is prepared for adhesive bonding, and the machined honeycomb core segments are applied. The segments are bonded together by an expanding adhesive splice during curing. The assembly is covered with an adhesive film again, and finally the lower skin prepreg layers are applied. The whole structure is cured in an autoclave, which is followed by debagging, CNC machining, NDT, hardware parts assembly, painting, and inspection steps. Fig. 5 Assembly process of the sandwich spoiler. B. Monolithic Spoiler I The monolithic spoiler presented in Sec. II.B is produced in two separate pieces, the upper skin and the backbone as indicated in Fig. 3. The geometry of the upper skin is simple in boundary and curvature. Therefore, the layup process can be automated easily. In the present case the laminate is produced from UD carbon fiber tapes, which are pre-impregnated with epoxy resin. The plies are applied automated with a tape layer machine. Then, the preproduced laminate is placed onto the layup mold, a vacuum bag is produced, and finally it is cured in an autoclave. All the complexity of the integrated stiffener design is concentrated in the backbone. Based on the geometry, dry carbon weaves with good drapability (e.g., five harness satin weaves) are applied to produce the preform, which is then infiltrated with resin and cured via RI. Its layup process is sketched in Fig. 6 and can be divided into four steps. In the first lamination step, full body plies, and reinforcement plies as needed are applied onto the cavity of the layup mold in defined orientations. Then, three tooling blocks are wrapped with carbon material (lamination step 2). Those subpreforms are assembled in lamination step 3 and are placed onto the already existing laminate to construct the CH. In lamination step 4 reinforcement and full body plies are placed to finalize the lamination. Fig. 6 Assembly process of carbon plies for the backbone of monolithic spoiler I. Before bond assembly of upper skin and backbone, both parts are checked by NDT, which is followed by cleaning and abrasion of the bonding surfaces. For bonding film adhesive is applied. The adhesive joint is cured in an oven; pressure is applied with a vacuum bag. Afterward the structure is finalized. Machining, NDT of the secondary bonding, dimensional inspection, hardware assembly, painting, and quality inspections are done. In Fig. 7 a complete manufactured spoiler is presented from different views. The center part of the upper skin is removed by intention to allow a look inside the hollow structure. Fig. 7 Views showing the manufactured monolithic spoiler I with partially removed upper skin. C. Monolithic Spoiler II In the second monolithic design proposal the spoiler body is produced in one layup and curing step as the main part, and C-spar components, which are simple CFRP parts, are produced to finally close the structure, as presented in Fig. 4. The spoiler body is produced with RI. Therefore, dry carbon fiber fabrics (UD and woven) are assembled to a preform. This process is described in Fig. 8. First, UD plies are applied onto the layup mold to form the upper skin surface. In a second step, all tooling blocks are wrapped with ply material. Here, fiber weaves with good drapability, for example, five harness satin weaves, are preferred. Those blocks are then placed onto the upper skin ply layup. In the fourth step the wedge is produced. It is running along the trailing edge at the tips of the subpreform blocks and is produced by stacking up of carbon ply material. Finally the assembly is topped with UD plies, which form the upper skin. Fig. 8 Assembly process of carbon plies for the body of monolithic spoiler II. For laminate design of the spoiler two options can be considered. Option I: The upper skin laminate and the lower skin laminate are produced from full body plies only. This would allow the use of standardized noncrimp fabrics only, and therefore a fast layup process is given. All doubler plies that are needed to reinforce the structure are integrated in the laminate of the subtooling blocks. Option I is presented in Fig. 8. Option II: As much doubler plies as possible are integrated in the lower skin and upper skin laminates. However, this would limit the use of noncrimp fabric. On contrary, the layup of the tooling blocks is kept simple and a decreased laminate thickness reduces the risk of skin wrinkling in the corner radii during resin infusion and curing. The C-spar sections are rather simple to produce, and a variety of semifinished fiber products and manufacturing processes can be applied efficiently. However, the same materials and resin infusion as for the spoiler body are considered in this publication. Of importance for bond assembly of spoiler body and C-spar is that the outer side of the C-spars is the molded surface. The reason is to provide a bonding surface of high geometric accuracy. Before bond assembly, NDT is performed on all subparts, which is followed by the preparation (cleaning and roughening) of the bonding surfaces. A film adhesive is used to join the surfaces. The adhesive joint is cured in an oven; pressure is applied with a proper clamping system. Afterward machining, NDT of the secondary bonding, dimensional inspection, hardware assembly, painting, and quality inspection are done. IV. Manufacturing Costs In the following, manufacturing costs are considered from the perspectives of used materials, applied processes, and the required equipment, for example, layup molds, machining jigs, and assembly jigs. The individual characteristics are compared with the sandwich spoiler. A. Cost Analysis Method Several software applications exist to perform cost estimations on composite applications in the aircraft industry. One common aspect of all software applications is to identify major cost contributors to the total manufacturing costs at early design phases of new development programs. It is known that costs allocated during the early design phases account for more than 80 of the final manufacturing cost of nonstandard parts. However, most of software applications are proprietary and only valid under certain restrictions. Further, benefits and limitations are listed. This application follows the parametric approach; knowledge gained by former development and production programs is collected and stored in a database. As long as a newly developed part is of sufficiently similar nature compared with the database, accurate cost estimation can be performed and a well-founded decision for the best design alternative is possible.