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Some of the methods used for experimental detection and examination of wake vortices are presented. The aim of the article is to provide the reader a brief overview of the available methods. The material is divided into two major sections, one dealing with methods used primarily in the laboratory, and the second part devoted to those used in field operations. Over one hundred articles are cited and briefly discussed.

As part of a general aviation airfoil development program being carried out under the direction of the NASA Langley Research Center, a 30% chord Fowler flap has been developed for the GA(W)-1 airfoil.. Wind tunnel tests at Wichita State University have demonstrated a c1max value of 3.80 for 40 deg flap deflection at a Reynolds number of 2.2 × 106. Effects of flap slot geometry have been systematically tested and optimum flap settings for any flight c1 have been obtained. Modification of the reflexed lower surface contour resulted in a reduced c1max with flap nested. Vortex generators provided an increase in c1max of 0.2 for flap nested and 40 deg flap along with a drag penalty at low c1 values. Flow visualization studies show that the stalling patterns for the new airfoil are characterized by an absence of leading edge separation for both the flap-nested and the 40 deg flap cases.

Metal cutting is a substantial constituent of airframe manufacturing. During the past several decades, it has evolved significantly. However, most of the changes and improvement were initiated by the machine tool industry and cutting tool industry, thus these new technologies is generally applicable to all industries. Among them, few are developed especially for the airframe manufacture. Therefore, the potential of high efficiency could not be fully explored. In order to deal with severe competition, the aerospace industry needs improvement with a focus on achieving low cost through high efficiency. The direction of research and development in parts machining must comply with lean manufacturing principles and must enhance competitiveness. This article is being forwarded to discuss the trend of new developments in the metal cutting of airframe parts. Primary driving forces of this movement, such as managers, scientists, and engineers, have provided significant influence to this trend.

This paper presents experimental methods for investigating large droplet impingement dynamics and for obtaining small and large water droplet impingement data. Droplet impingement visualization experiments conducted in the Goodrich Icing Wind Tunnel with a 21-in chord NACA 0012 airfoil demonstrated considerable droplet splashing during impingement. The tests were performed for speeds in the range 50 to 175 mph and with cloud median volumetric diameters in the range of 11 to 270 microns. Extensive large droplet impingement tests were conducted at the NASA Glenn Icing Research Tunnel (IRT). Impingement data were obtained for a range of airfoil sections including three 36-inch chord airfoils (MS(1)-0317, GLC-305, and NACA 652-415), a 57-inch chord Twin Otter horizontal tail section and 22.5-minute and 45-minute LEWICE glaze ice shapes for the Twin Otter tail section. Small droplet impingement tests were also conducted for selected test models.

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.

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.

During this study, a number of 8.5-inch by 11.5-inch flat honeycomb sandwich panels were inflicted with low energy impact damage, inspected non-destructively, and tested for residual in-plane compressive strength. Each panel had either a 3/8-inch or 3/4-inch low density Nomex honeycomb core, and either 2-ply, 4-ply or 6-ply face sheets. The face sheets were either carbon or Eglass (prepreg) fabric. The panels were either clamped or simply supported in a test fixture during impact from a gravity assisted drop mechanism, and impacted with either a 1-inch or 3-inch diameter spherical indenter. After impact the damage to each panel was characterized by (1) ultrasonic through-transmission to obtain a c-scan representing planar damage area, (2) indentation volume and depth, and finally (3) visual inspection to rate the damage according to a predetermined rating scale. The panels were then tested for in-plane compressive strength.

A unique cascade test facility has been developed for use in the Wichita State University (WSU) water table. Although small in scale, the WSU water table has the advantage of low cost and the ease with which test conditions can be varied. Water table facilities have been used in the past for cascade experiments, especially as analogies for compressible flow visualization of turbine cascades. However, the lack of a quantitative measurement technique at low speeds has precluded the use of the water table as an analogy for testing subsonic compressors and turbines. In the present experiment, the hydrogen bubble flow visualization technique is used to generate bubble time lines, and a CCD (Charge Coupled Device) video camera system captures and digitizes these time line images. A VisualBASIC® computer program is then used to determine the wake velocity profile based on the difference in bubble line positions at successive intervals of time.

The work presented in this paper is part of a long-term research program to explore methods for improving bleed air system performance. Another objective of this research is to provide detailed experimental data for the development and validation of simulation tools used in the design and analysis of bleed air systems. A business jet wing was equipped with an inner-liner hot air ice protection system and was extensively instrumented for documenting system thermal performance. The wing was tested at the NASA Glenn Icing Research Tunnel (IRT) for representative in-flight icing conditions. Data obtained include bleed air supply and exhaust flow properties, wing leading edge skin temperatures, temperatures and pressures in the interior passages of the bleed air system, flow properties inside the piccolo tube, photos of run back ice shapes and ice shape traces. Selected experimental results for a warm hold icing condition are presented in this paper.

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.

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.

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.

In this research helical Carbon Nanotubes (CNTs) with various weight percentages as an additional reinforcement were used. The objective was to investigate the effectiveness of helical geometries of the CNTs to form interlocking mechanisms with the resin and the traditional microfiber reinforcements to improve the overall performance of the composite structures and assemblies. In this study, ASTM D2344/2344M-16 is used to study the short beam strength of the laminated nanocomposites and evaluate the benefit of the mechanically interlocked helical CNTs reinforcement. Overall, three sets of composite laminates (i.e., with neat epoxy, and with two different wt% of Helical CNTs reinforced epoxy) were fabricated per ASTM standard D2344/2344M-16. Adequate test specimens were prepared and then they were tested per ASTM standard. The test results were analyzed and evaluated to determine the effects of helical CNTs on short beam strength of the laminated nanocomposites.

A methodology for developing spray rigs for icing cloud simulation is presented. This methodology includes Computational Fluid Dynamics (CFD) analysis and icing tunnel experiments and was applied to design a spray rig system for a small airborne icing tanker. An in-house spray system was developed and tested in a laboratory to assess two commercially available nozzles - a single-jet type and a multi-jet type - which were capable of producing both FAR Part 25 Appendix C and SLD icing clouds. Spray rig characteristics evaluated during the laboratory tests included air and water flow rates as well as droplet size and distributions. The effects of airspeed and nozzle spacing on spray plume size and uniformity were investigated in a small icing tunnel facility with a two-nozzle spray rig. The experimental data were compared with three-dimensional numerical simulation results obtained with the FLUENT software.

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.

Ice particles shed from aircraft surfaces are a safety concern because they can damage aft-mounted engines and other aircraft components. Ice shedding is a random and complex phenomenon. The randomness of the ice fragment geometry, size, orientation and shed location in addition to potential particle breakup during flight poses considerable simulation challenges. Current ice shedding analysis tools have limited capabilities due to the lack of experimental aerodynamic coefficients for the forces and moments acting on the ice fragment. A methodology for simulating the shedding of large ice particles from aircraft surfaces was developed at Wichita State University. This methodology combines experimental aerodynamic characteristics of ice fragments, computational fluid dynamics, trajectory analysis and the Monte Carlo method to provide probability maps of shed particle footprints at desired locations.

The necessity of avoiding the destructive and non-repeatable FSST (Full Scale Sled Test) makes it desirable to devise a cheaper and more repeatable method which can supplant this test procedure. This need developed the HCTD (HIC Component Testing Device) which is capable of providing conservative HIC results with higher repeatability. The computational model of the HCTD is validated against one of the tests conducted at CAMI with polyethylene foam. This validated model is used to conduct a series of tests with input parameters similar to the sled test to develop the correlation between the sled test and HCTD. This study hence concludes that a validated computational model of HCTD can be successfully utilized to address the HIC compliance issues for a foam padded surface.

Global warming concerns are stimulating accelerated research and development of alternative fuels and propulsion systems for automobiles. The potential application of these emerging technologies to airplanes is reviewed. Preliminary designs of zero greenhouse gas emission airplanes using hydrogen fuel and either internal combustion or fuel cell-electric motor propulsion are presented for a wide body jet transport, medium jet transport, business jet, and single engine propeller airplane. The hydrogen fueled internal combustion engine airplanes offer the easiest path to zero emissions, but the greater efficiency of the fuel cell airplanes allows designs requiring substantially less fuel. The single engine propeller airplane is the easiest to modify for hydrogen fuel, because of the relatively high mass and volume of the engine being replaced. Technology improvements needed to make zero emission airplanes viable are suggested.

Research has been conducted on the nature of airfoil behavior at pre- and post-separated angles of attack. Detailed wind tunnel studies have been made of boundary layer and wake fields for the GA(W)-1 airfoil, and the airfoil with a 0.3 chord Fowler flap. Experimental data are compared with theoretical predictions from a multi-element viscous flow computer program. Theoretical predictions are reasonably accurate for unseparated flows, but have serious errors when separation is present. Some recent techniques for modeling post-separated flow behavior are discussed in light of the present experiments.