Search Results

Most composite assemblies and structures generally fail due to weak interlaminar properties and poor performance of their bonded joints that are assembled together with an adhesive layer. Adhesive failure and cohesive failure are among the most commonly observed failure modes in composite bonded joint assemblies. These failure modes occur due to the lack of reinforcement within the adhesive layer in transverse direction. In addition, the laminated composites fail due to the same reason that is the lack of reinforcement through the thickness direction between the laminae. The overall performance of any composite structures and assemblies largely depends on the interlaminar properties and the performance of its bonded joints. Various techniques and processes were developed in recent years to improve mechanical performance of the composite structures and assemblies, one of which includes the use of nanoscale reinforcements in between the laminae and within the adhesive layer.

The Refill Friction Spot Joining (RFSJ) is an emerging solid-state spot welding technology that thermo-mechanically creates a molecular-level bond between the work-pieces. RFSJ does not consume any filler or foreign materials so that no additional weight is introduced to the assembly. As the solid-to-liquid phase transition is not involved in RFSJ in general, there is no lack of fusion or material deterioration caused by liquefaction and solidification. Unlike the conventional friction stir spot welding, RFSJ produces a spot joint with a perfectly flush surface finish without a key or exit hole. Currently, the aerospace industry employs solid rivets for fastening the primary structures as they meet the baseline requirements and have well-established standards and specifications.

Presented here are analyses and statistical summaries of data collected from 11,299 flight operations recorded on 6 BE-1900D aircraft during routine commuter service over a period of three years. Basic flight parameters such as airspeed, altitude, flight duration, etc. are shown in a form that allows easy comparison with the manufacturer's design criteria. Lateral ground loads are presented for ground operations. Primary emphasis is placed on aircraft usage and flight loads. Maneuver and gust loads are presented for different flight phases and for different altitude bands. In addition, derived gust velocities and various coincident flight events are shown and compared with published operational limits.

Dynamic aircraft seat regulations are identified in the Code of Federal Regulations (CFR), 14 CFR Parts § 23.562 [1] and § 25.562 [2] for crashworthy evaluation of a seat in dynamic environment. The regulations specify full-scale dynamic testing on production seats. The dynamic tests are designed to demonstrate the structural integrity of the seat to withstand an emergency landing event and occupant safety. SAE standard AS 8049 [3] supports detailed information on dynamic seat testing procedure and acceptance criteria. Full-scale dynamic testing in support of certification is expensive and repeated testing due to failure drastically increases the expense. Involvement of impact environment, flexibility in interior configuration and complicated nature of seat engineering design makes this problem quite complex, so that classical hand calculations are practically impossible.

The fatigue behavior of Hilok fastener joints under constant amplitude loading has been investigated experimentally. The effects of load transfer in an unbalanced joint configuration was characterized in terms of a stress severity factor relative to the open-hole configuration. The experimental data indicates that the clamp-up forces dominate the performance of fastener joints with the open-hole fatigue life being the lower bound at the stress levels investigated. The failure modes were observed to transition from a net-section type failure across the minimum section to a fretting induced failure at some distance from the hole. The experimental data has been used to develop stress severity factors to be used as a measure of the fatigue quality of the fastener joints.

A supplier satisfaction survey was developed and administered to 129 Aircraft suppliers who are AS9100 registered. The primary objective of the survey was to assess organizational benefits, attributed to the AS9100 standard, and registration process difficulties. Survey results from 49 responses indicated that the primary reason for seeking AS9100 registration was customer requirement, followed by improving production and service. Further analysis indicated that the top three difficulties were evaluating effectiveness of employee training, obtaining and analyzing data on customer feedback and satisfaction, and monitoring and measuring processes. The top three reported benefits, improved quality awareness among employees, an increase in employee training, and improved internal communication, respectively, were all non-financial in nature.

A series of carefully selected tests were used to isolate the coupled influence of various combinations of the number of facesheet plies, impact energies, and impactor diameters on the damage formation and residual strength degradation of sandwich composites due to normal impact. The diameter of the planar damage area associated with Through Transmission Ultrasonic C-scan and the compression after impact measurements were used to describe the extent of the internal damage and residual strength degradation of test panels, respectively. Standard analysis of variance techniques were used to assess the significance of the regression models, individual terms, and the model lack-of-fit. In addition, the inherent variability associated with given types of experimental measurements was evaluated.

This study presents the prediction of the dimensional variation of holes due to sheet metal bending using the hydroforming technique. Sheet metal with pre-drilled holes was evaluated for a bending operation using a hydroforming technique. Sheet metal with a variety of thicknesses, bending radii, and bending angles was evaluated. Variation in the dimensional tolerance due to the bending was attained using the minimum radial separation method. A dataset of dimensional variation in the holes was developed and used for development of the artificial neural network, which was able to predict the dimensional variation of the hole if an unknown pattern of inputs was provided.

The coupled influence of material configuration (number of facesheet plies, core density, core thickness) and impact parameters (impact velocity and energy, impactor diameter) on the impact damage resistance characteristics of sandwich composites comprised of carbon-epoxy woven fabric facesheets and Nomex honeycomb cores was investigated using empirically based quadratic response surfaces. The diameter of the planar damage area associated with TTU C-scan measurements and the peak residual facesheet indentation depth were used to describe the extent of internal and detectable surface damage, respectively. Estimates of the size of the planar damage region correlated reasonably well with experimentally determined values. For a fixed set of impact parameters, estimates of the planar damage size and residual facesheet indentation suggest that impact damage development is highly material and lay-up configuration dependent.

An aging aircraft accumulates fatigue cracks commonly referred to as multiple site damage (MSD). A simplified engineering fracture mechanics model, generally referred to as the linkup model (or plastic zone touch model), has been used with some success to describe the MSD fracture phenomenon in 2024-T3 aluminum panels. A disadvantage of the linkup model is that it gives excessively inaccurate results for some configurations. A modified linkup model has been developed through empirical analysis of test data taken from unstiffened panels with MSD cracks at open holes. The modified linkup model was then validated with test data from stiffened panels including single-bay panels with the lead crack centered between stiffeners and two-bay panels with the lead crack centered beneath a severed stiffener. Further validation of the modified linkup model was done with test data from panels with bolted lap joints. Test results were obtained from 112 different panels.

This paper presents the experimental study of hole quality parameters in the drilling of titanium alloy (6Al-4V). Titanium alloy plates were drilled dry using three types of solid carbide drills i.e. 2-flute helical twist drill, straight flute and three-flute drill. The objective was to study the effects of process parameters like feed rate, speed and drill bit geometry on the hole quality features. Typical hole quality features in a drilling process are the hole quality measures such as surface roughness, hole diameter, hole roundness and burr height. The results indicate that proper selection of speed, feed rate, and drill geometry can optimize metal removal rate and hole quality.

The competitive market has forced the industry to develop methodologies to reduce lead-time of the products without sacrificing quality. One of the major metal removal operations in the aerospace industries is drilling. Over 100,000 holes are made for a small single engine aircraft. Naturally, demand for faster production rate results in the demand for high-speed drilling. But the cost of hole-making operations becomes a significant portion of the total manufacturing cost. This paper discusses the high speed drilling of Al-2024-T3 alloy, the effect of feed and speed on hole quality features like oversize, roundness error, burr height and surface roughness.

The residual strength of an aluminum panel with a centric hole and one cracked ligament was investigated experimentally. Each of the 7075-T6 aluminum panels which were tested included a cracked ligament of varying length on one side of the centric hole and an uncracked ligament on the other side of the hole. The failure of such a panel subjected to uniform tensile loading normally occurs according to the lower of two modes: brittle fracture or a net section type of yielding. On the other hand, the question of whether one or both ligaments fail is not easily answered. Results show that one or two ligament failure depends upon test conditions such as crack length and loading method. For short crack lengths, the uncracked ligament will fail almost simultaneously with the failure of the cracked ligament.

Multiple site damage (MSD) on aging aircraft accumulates from fatigue loading over a period of time. For ductile materials such as 2024-T3 aluminum, MSD may lower the strength below that which is predicted by conventional fracture mechanics. An analytical model referred to as the linkup (or plastic zone touch) model has previously been used to describe this phenomenon. However, the linkup model has been shown to produce inaccurate results for many configurations. This paper describes several modifications of the linkup model developed from empirical analyses. These modified linkup models have been shown to produce accurate results over a wide range of configurations for both unstiffened and stiffened flat 2024-T3 panels with MSD at open holes. These modified models are easy to use and give quick and accurate results over a large range of parameters.

An experimental study was conducted to investigate ice-adhesion on clean and coated aluminum surfaces. A test apparatus using the parallel plate linear shear technique was designed along with a data acquisition system for conducting the tests and recording the experimental data. A low pulling rate was applied to specially prepared test specimens for measuring the strength of ice adhesion for a range of test conditions. The effects of surface roughness, surface contamination, and water impurity on ice adhesion were investigated. In addition, tests were conducted to evaluate the effectiveness of a low ice-adhesion coating applied to aluminum test specimens. The results obtained showed that the bond between ice and metal was considerably lower for tap water than for distilled water. For the clean and coated aluminum surfaces the strength of ice adhesion varied with specimen roughness. However, no clear trend was established between ice adhesion strength and surface roughness.

Drilling is one of the most widely applied manufacturing operations. Millions of holes are drilled today in manufacturing industries especially in aerospace industry where high quality holes are essential. Rejection and rework rate of the products because of the bad hole is quite high. In this research graphite/honeycomb composite material and aluminum sheet metal has been used. The results show that drill geometry, speed and feed rate have substantial effects on the hole quality and also there was gradual variation of the thrust and lateral forces with feed rates.

Flight test experiments were conducted to measure the extent and nature of natural laminar flow on a smoothed test region of a swept-wing business jet wing. Surface hot film aneraometry and sublimating chemicals were used for transition detection. Surface pressure distributions were measured using pressure belts. Engine noise was monitored by a microphone attached to the wing surface to study possible acoustic effects on stability of the laminar boundary layer, Side-slip conditions were flown to simulate changes in effective wing sweep. Flight instrumentation and ground data analysis techniques and a method for measuring intermittency of turbulence are described, Correlation was obtained between the hot film gage signals and chemicals for transition detection. Cross-flow vortices were observed for some flight conditions. Results of spectral and statistical analysis of the hot film signals for various flight test conditions are presented.

Laminated metal-metal composites can have attractive fracture toughness properties; they also offer potentially good fatigue performance. These attributes are reviewed and prospects for improvement discussed. Weak interlaminar bonds are seen to be important, while quite thin layers seem to be most promising for laminates of higher strength materials. The experimental program utilized 0.033 in (0.84 mm) thick laminae of 7075-T6 aluminum alloy, adhesively bonded. Eight-layer composites were compared with solid sheets of nearly the same total metal thickness. Both fracture toughness and fatigue properties were determined. Kc values of more than double the KIc for this alloy were observed in the laminates, while fatigue performance as indicated by comparative S-N curves was found to be slightly improved.