Kinematic equations were developed to describe the dynamic motions of the aircraft piston engine components in terms of time dependent position, velocity, and acceleration relationships. Using the Aeronca E-113 engine as a case study, the brake mean effective pressure (BMEP) rating was used to model the cylinder gas pressure profile. The moments of inertia of the dynamic components including connecting rod, crankshaft, and propeller were measured using a pendulum swing method. Representative values were obtained for inertial and gas pressure forces acting on crankshaft journals, connecting rods, and cylinder walls. The resulting model can help in the design of crankshafts and other dynamically loaded parts to resist failure due to fatigue.

Digitization of physical parameters for the display and recording by computers is the essential aspect of any airborne data acquisition system. The objective of this thesis was to develop a data acquisition system for General Aviation research and certification flight testing based on a low-cost Commercial Off-The- Shelf (COTS) hardware, in particular, a common glass cockpit system for experimental aircraft. A kneeboard computer was used to monitor data communications between the various devices of the Grand Rapids Technology (GRT) Electronic Flight Information System (EFIS). The monitored data was then displayed for use in-flight, and recorded aboard the aircraft for post-flight data reduction. The developed system and software was tested in simulation on virtual and actual hardware, on an Extra 300 in ground testing, and in flight. An in flight air-data calibration and several common stability and control certification test points were flown to evaluate and demonstrate the usefulness of the system. Special consideration was paid to work flow prior to, during, and after the flight with the overall goal of reducing the time required for data reduction. The output of this research work includes software for decoding data files logged on one common low-cost EFIS, software for monitoring, displaying, and recording EFIS data on a kneeboard computer in-flight, and tools for managing and viewing data files after the flight. From this research work, it is concluded that commercially available EFIS systems do in fact provide a core data set which is useful in flight research and flight test certification programs. The 15 Hz sampling rate of the GRT system was more than sufficient for all the test points evaluated as a part of this research. The cost of the tested hardware was less than $10,000 at current pricing (2009). The resultant system is compact, adds little weight to a test aircraft, has few interfaces to aircraft systems, and allows for future growth and the incorporation of new sensor types and interfaces. The addition of a flight test air-data boom with angle of attack and sideslip vanes and control position and force sensors would create a very complete data acquisition package without the expense of purpose designed hardware.

A study was undertaken to assess the effectiveness of the CF188 main landing gear upgrade on reducing Planing Link Mechanism failures. Two main landing gear configurations were studied: the prototype configuration and, for comparison purposes, the current configuration referred to as the baseline. Under this study, the flight test data that was analyzed came from key measurements recorded during maintenance rigging procedures, pilot ground handling quality ratings, and from over 80 landings at different descent rates and aircraft attitudes. Landings consisted of touch and go, full stop, cable overrun and cable engagement. The aircraft that was used through the flight test program had both its main landing gears instrumented. While the prototype configuration had minimal impact on the ground handling characteristics, it demonstrated promising results during maintenance activities as well as loads distribution during landings. The prototype was easier to rig which will allow the use of tighter limits further standardizing the complex maintenance procedure. From both a static and dynamic point of view, the hold down force was significantly increased. Within the scope of this study, it was found that the CF188 main landing gear upgrade will reduce Planing Link Mechanism failures.

The purpose of this thesis was to investigate if it would be possible to integrate a developmental Radar Warning Receiver named “The Puffer” onto a Tier 2 Unmanned Aerial Vehicle, and incorporate the Multifunctional Information Display System/Low Volume Terminal into the UAV‟s ground control station. This integration of systems would become a low cost platform that could provide an Electronic Surveillance and geolocation capability of known mobile threat systems. The results of this investigation showed that the Puffer could be integrated on to the Tier 2 UAV with minor modifications. To control and fully integrate the downlink messages from multiple UAVs plus add the capability to send the information out to other units over Link16 would require a major hardware effort with a sizable software integration effort. While this would be an extensive project, the results could be done at a significant cost saving compared to the manned platforms in use today.

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.

To complement previous studies on the bidirectional vortex, we attempt to characterize viscous effects in both the axial and radial directions along the sidewall with standard asymptotic techniques. The actual boundary layer present in the chamber will be a composite of both axial and tangential shearing layers. Since the tangential velocity is completely dominant, we expect the defining characteristics of the composite boundary layer to be congruent with the tangential layer in both thickness and axial invariance. The analytic analysis is outlined first by, the formulation of the boundary layer equations via Prandtl’s method. Next, asymptotic techniques are applied to linearize and rigorously truncate the governing equations from PDEs to more manageable ODEs. A scaling transformation is applied to resolve the rapid changes near the wall. Due to the nature of the outer solution, a dependent variable transformation is applied to recover constant boundary conditions. The viscous corrections are matched to the outer solution via Prandtl’s matching principle. We see a similar form in all three wall corrections; the axial and radial presented here and the swirl previously formulated. This can be expected to some extent because of the similarity of the asymptotic assumptions and the linearization techniques used in all three cases. Although the assumption that curvature terms can be neglected is never made, they are found to be asymtotically small and the problem then parallels the case of a one-dimensional Cartesian boundary layer. It can be seen that all viscous corrections along the wall are strongly dependent on the value of the vortex Reynolds number, V. This parameter shows up naturally in all vector directions. With the new corrected solutions, other key features of the flowfield can be revisited.

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.

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.

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.