Two of the most recognizable competitors in this new era of Euro-American competition are Airbus and Boeing. These two competitors are the world’s leading aircraft manufacturers and their competition is shaping the face of the commercial aviation industry. The competition between these two companies is for the “World’s Leading Commercial Airline Manufacturer.” This title is currently held by Airbus who has lead yearly production and taken more orders over the past three years. Boeing has fallen behind in being the technological leader due to a falling budget, poor industrial model, and ethical practices. As a result of losing ground to Airbus over the past few years, Boeing has been continuously restructuring itself in an effort to compete in this new era of competition. By comparing these two companies the author’s objective was to determine a clear path forward for Boeing. To make this determination the author examined both companies while making a detailed analysis of certain areas. Three focus areas were chosen based on initial analysis and the magnitude of their effects. These focus areas included subsidies, technology integration, and vision of the future. Airbus’ early successes can be attributed to an unfair subsidy advantage that is now at the center of this competition. The next aspect at the center of this competition is their philosophies on automation and the implementation of technology. Boeing and Airbus have different philosophies on the implementation of automation. While both philosophies are sound, there are advantages and disadvantages to each. Finally, each company in an attempt to gain an edge in the competition has staked its future on what each believes to be the future direction of commercial aviation. Airbus has gone with the philosophy of the jumbo jet with the A380, whereas Boeing has opted for targeting the medium range market with the 787. The analysis of these two companies shows their difference in philosophies with regard to embracing new technologies in aircraft design and manufacturing. This thesis examines how Airbus has risen as the “World Leading Commercial Aircraft Manufacturer” and Boeing’s need to self-optimize.

Despite the recent improvements to the T-38 airframe and engines, the United States Air Force is still seeking ways to improve the aircraft’s takeoff, cruise, and landing performance. One potential way to improve the performance is to change the design of the wing. Using the Digital Performance Simulation aircraft-performance computer code, a T-38C performance evaluation sensitivity study was performed by parametrically varying the wing design. The computer model was a three degree of freedom, pointmass, batch simulation. The design changes investigated included varying aspect ratio with constant wing area, varying wing area with constant aspect ratio, and the addition of a winglet. These preliminary design estimates compared the differences in takeoff, cruise, and landing phases resulting from the modifications to the current baseline configuration. Using a variety of aerodynamic theories, new aircraft lift curves and drag polars were developed. These new aerodynamic models were then used in the computer simulation to determine the new aircraft performance during the various phases of flight. While incremental improvements were made in maximum range, maximum speed, and landing distances, a major improvement in the single-engine climb performance was found with a small increase in wing area from the baseline value of 170.0 square feet to 183.7 square feet. With a weight gain of only approximately 138 pounds, the operational envelope of the aircraft can be significantly increased. This larger wing will provide a 10 knot improvement in single engine takeoff speed and a 7.5% reduction in landing distance and will allow continued operation of the aircraft in the most demanding environmental conditions.

Over the past decade the United States Army has used a Full-Authority Digital Electronic Control (FADEC) system to control fuel flow to the engine of the OH-58D helicopter. Currently, part of the training is primarily conducted for the scenario of a FADEC system failure in the aircraft. Because of the complexity of this task, a number of accidents have occurred resulting in minor to severe damage to the aircraft. The United States Army has recently fielded two OH-58D Operational Flight Training Simulators in an effort to increase training efficiency and effectiveness. It is anticipated that the simulators will provide a safer environment and an effected transfer of training to the aircraft. Currently the OH-58D training unit has implemented the simulator into the manual throttle stage of training. This implementation has occurred through verification and validation of the Program of Instruction (POI) currently in use. An investigation into the transfer of training from the simulator to the aircraft was conducted to further optimize the distributions of training time in the simulator versus the aircraft. The primary source of data was collected from aircraft and simulator trials and flight hours to evaluate the transfer effectiveness ratio. The secondary source of data was collected through the use of pilot surveys and questionnaires. The pilots reported a mean workload rating of 2.52 using the Bedford Workload Rating Scale in the aircraft after the simulator, which indicates a low workload. The Pilots reported mild to moderate simulator sickness symptoms after flying in the simulator. A total severity score of 20.06 was computed through the use of the Simulator Sickness Questionnaire. When compared to other helicopter simulators this score is fairly high. Overall there were low Pilot-Vehicle Interface problems in the simulator and aircraft. There was no decline in Situational Awareness from the simulator to the aircraft. The overall Transfer Effectiveness Ratios indicated a positive Transfer of Training. The current Program of Instruction and simulator hours are validated. The focus in the simulator should be placed on Method of Instructions step two “failure at a hover” and step four “running landing or approach to a hover.”

In this article, we consider the vorticoacoustic flowfield arising in a rightcylindrical porous chamber with uniform sidewall injection. Such configuration is often used to simulate the internal gaseous environment of a solid rocket motor (SRM). Assuming closed-closed acoustic conditions at both fore and aft ends of the domain, the introduction of small disturbances in the mean flow give rise to an axially traveling vortico-acoustically dominated wave structure that our study attempts to elucidate. Although this problem has been formulated before, it is reconsidered here in the context of WKB perturbation expansions in the reciprocal of the crossflow Reynolds number. This enables us to uncover multiple distinguished limits along with new asymptotic solutions that are presented for the first time. Among them are WKB approximations of type II and III that are systematically evaluated and discussed. The WKB solutions are shown to exhibit a peculiar singularity that warrants the use of matched asymptotic expansions to produce uniformly valid representations. Our solutions are obtained for any characteristic mean flow function satisfying Berman’s similarity condition for porous tubes. They are also derived to an arbitrary level of precision using a recursive formulation that can reproduce each of the asymptotic solutions to any prescribed order. Finally, our solutions are verified numerically over a wide range of physical parameters and through limiting process approximations.

This work investigates the optimal topology of an actively controlled piezoelectric actuator bonded to an elastic cantilever beam under steady-state harmonic loading near the first natural frequency of the beam. The actuator is discretized using finite elements, and control is applied to the actuator based on the sensor’s degrees of freedom using proportional control. This study investigates the optimal distribution of actuator material for one and five layers of finite elements. The optimized actuator topology shows substantial improvement over initial piezoelectric topologies and over traditional actuator placement. This study has two main topics, material homogenization and topology optimization. The piezoelectric actuator is homogenized to determines it’s relation to volume fraction of material. This continuous relationship to volume fraction is a more “realistic” material variation compared with the typical artificial material model used in topology optimization. The actuator topology of the actuator is optimized to minimize the vibration amplitude. To the author’s knowledge, little work has focused on the optimal topologies of a piezoelectric actuator to minimize the amplitude displacement at the first mode of a cantilever beam.

The purpose of this study was to determine the feasibility of applying the Maintenance Error Decision Aid (MEDA) that was developed by Boeing to general aviation maintenance shops, either in its current form or with limited modification. The MEDA investigation process has been implemented successfully by several major airlines and it was assumed that general aviation could also benefit from this safety enhancing process. Because of the nature of the MEDA process, this paper only addresses the feasibility of applying the MEDA process to large shops. After consulting aviation professionals and performing extensive research, a questionnaire was created and sent to numerous general aviation (GA) maintenance managers to determine their opinion of the feasibility of the application of MEDA to GA. A total of 6 responses were received and analyzed, from which it was concluded that the MEDA system could enhance safety in general aviation with certain alterations to the system.

This thesis investigates the effects of personnel wind deflector devices on the level flight performance of an MD-500D helicopter configured with external passenger provisions. Numerous helicopter organizations operate with external passenger configurations. These configurations result in personnel exposure to high winds and an increase in parasite drag. Level flight performance is degraded by the increase in parasite drag caused by the external passengers. Wind deflectors were mounted on the forward portion of the fuselage to protect external passengers from the effects of wind exposure (high wind loads and wind chill factor) by deflecting the wind away from the fuselage. The purpose of this investigation is to determine the effects of the wind deflector modification on level flight performance; specifically the change in: engine shaft horsepower required, equivalent flat plate area, maximum attainable endurance, and maximum attainable range. Four helicopter external configurations were test flown, and the data compared to determine the affects on performance caused by the wind deflector modification. The constant W/σ flight test technique was used in measuring the power required for level flight in each of the four configurations. With four manikins mounted outside the aircraft and wind deflectors installed, the maximum level flight speed and maximum range increased by 4.8% and 7.1% respectively. These percentages are relative to the aircraft with four manikins mounted outside the aircraft and no wind deflectors installed. Without manikins mounted outside the aircraft and wind deflectors installed, the maximum level flight speed and maximum range decreased by 7.6% and 11% respectively. These percentages are relative to the aircraft without manikins or deflectors mounted outside the aircraft. Maximum endurance was not affected by the wind deflector modification.

The Rijke tube device has been employed since its invention in 1859 in the experimental study of many examples of thermo-acoustic phenomena. The device exhibits generation of acoustic oscillations by heat energy supplied to the flow field in the fashion of a selfexcited oscillator. In recent times, the Rijke tube has proved to be a valuable tool in simulation of combustion instability phenomena in rockets and industrial burners. Despite the simplicity of the device, the Rijke tube simulates most important geometrical and physical features that lead to the growth of nonlinear pressure oscillations in combustion chambers. For example it provides a through-flow as in a rocket chamber and is fixed with an energy source that can cause unsteady combustion. The open ends and geometrical simplicity leads to easy accessibility for instrumentation to make measurements that would not be possible in actual combustion chambers. During operation, wave motions are generated by transfer of energy from a heated grid placed at a point within the chamber that can be related to theoretical models for the phenomenon by Rayleigh and other investigators. However, initially, there is exponential growth of these oscillations to high amplitude and transition to a nonlinear limit cycle at a nearly fixed amplitude (usually lasting several seconds) due to natural nonlinearities in the system. The hypothesis advanced in this thesis to explain this nonlinear limiting effect that is the wave steepening occurs in a manner analogous to similar generation of steep wave fronts in rocket motor chambers. The latter proposal is based on: 1) direct observation (using Schlieren techniques) of traveling shock-like waves in axial mode instability, 2) correlation of the observed waves with spectral components similar to that of sawtooth structure, and 3) theoretical calculations showing that the limit amplitude phenomenon is directly related to the cascade of energy from lower frequency standing acoustic modes to higher harmonics leading to characteristic spectrum similar to that of a traveling steep-fronted wave. In prior research, the ‘mechanism’ of initiation of instability in the system has been the main focus. The goal of the research described in this thesis is to measure and to characterize the signal produced during the high- amplitude (nearly steady state) oscillations at the limit cycle. The intent was to demonstrate in a very simple way that the gas motions produced during the limit cycle in the Rijke tube have the same characteristics observed in many years of rocket testing. The observations again verify the great utility of the Rijke tube in seeking better understanding of the analogous rocket instability.

The purpose of this study was to evaluate and provide recommendations for optimizing the Pilot Vehicle Interface for components of data link systems employed on and currently in design for the F/A-18 Hornet and Super Hornet. Data was gathered using human factors research methodologies including descriptive studies, experimental research, and evaluation research. Additionally, flight and lab tests were used to gather data on systems that were mature enough in development. Overall, the study revealed that the interfaces for the systems evaluated could be modified in order to provide more situational awareness to the operator, allow for more logical display of information, and improve the operator interface with the overall effect of increasing the efficiency of the weapon system as a whole. While hardware display improvements would solve many display limitation problems with the Situational Awareness format, there are potential software solutions that were assessed to be adequate and much more cost effective. The software solutions will aid in displaying, on the Situational Awareness and expanded formats, information that is currently omitted under certain conditions. Decluttering the Track Number search format and Helmet Mounted Display while displaying pertinent information in a more concise manner will increase the efficiency with which the operator processes it. Displaying information on the Close Air Support format in a more usable format with the appropriate level of detail will help reduce the potential for fratricide. Standardizing the push button labels associated with the “cease” command function on the RECALL and NETS formats will significantly reduce operator workload, errors, and required training.

A limited aerodynamic, stability and control, and task evaluation of a new rocket-powered Mars airplane design was conducted. The Mars airplane design, designated the Argo VII, was patterned after the NASA ARES-2 design. The aerodynamic and stability and control parameters of the Argo VII were determined using analytical and computational techniques and were comparable to those of the ARES-2. The Argo VII was predicted to be statically stable and damped in all axes on Earth and Mars. A series of flight tests were performed using a MATLAB Simulink-based flight simulation program to assess the performance, longitudinal flying qualities, and mission effectiveness of the Argo VII flying on Earth and Mars. At an assumed Mars mission flight condition of 2 km (6,562 ft) altitude and 0.65 Mach, the Argo VII had a maximum range lift coefficient of 0.44, a maximum lift-todrag ratio of 15.5, and a maximum endurance lift coefficient of 0.76. The Argo VII was dynamically stable and damped in the longitudinal axis. At the Mars mission flight condition, the long period had a damping ratio of 0.04, damped and undamped natural frequencies of 0.0423 rad/s (2.42 deg/s), and time to half of 409.6 sec. The short period had a damping ratio of 0.2, damped natural frequency of 7.39 rad/s (723 deg/s), undamped natural frequency of 7.54 rad/s (432 deg/s), and time to half of 0.46 sec. At the Mars mission flight condition, the aircraft had a specific excess power of 5.8 m/s (19.02 ft/s). At all Mars altitudes evaluated, the fastest way for the aircraft to change altitudes was to climb to the desired altitude at a constant equivalent airspeed. Mars mission aircraft task evaluations were performed using Mars simulation scenery to validate the predicted aircraft range and climb and descent performance. The aircraft range evaluation resulted in an aircraft maximum range of 373 km (232 mi). The predicted aircraft maximum range was 500 km (311 mi). The climb and descent evaluations resulted in aircraft performance that was similar to the predicted aircraft performance. This research illustrated that the Argo VII Mars aircraft design can provide a viable means of acquiring scientific data on Mars.