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.

Subsonic cavity flow tests of an L/D = 3.5 cavity, with three different diameter rods in crossflow, 1/8", 3/16", and 1/4", were conducted using the High Speed Wind Tunnel in the University of Tennessee Space Institute’s Gas Dynamic Laboratory. The average Mach number flow over the duration of the four phase testing sequence was 0.52, with a unit Reynolds number of 13.8 x 106. With the use of a dynamic pressure transducer and a laser PIV system, Spectral and Flow Visualization data was collected with aim of investigating the effect of the rods in crossflow on cavity flow. However, for reasons beyond the control of this investigation, a converging-diverging supersonic nozzle was used in place of a subsonic nozzle. As a result, the separated, or near separated, flow on the diverging side of the nozzle created a region of low kinetic energy flow approximately 5 mm above the floor of the tunnel test section. Despite the presence of this undesirable feature, the Baseline cavity, without a rod in crossflow, was found to resonate at 1413 Hz and produced an average peak amplitude tone of 148.7 dB SPL. The effect of placing different diameter rods in the crossflow was to reduce the amount, and intensity, of shear layer interactions, by helping to loft the flow over the trailing edge of the cavity. The best results were achieved with a 1/4" diameter rod, which, on average, provided 15.1 dB SPL of acoustic suppression. It was concluded that the suppression observed in this particular experiment was the result of blockage and lofting effects, which helped the shear layer to span the length of the cavity and reduce the intensity of the shear layer interactions at the trailing edge.

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.

In this thesis, a semi-analytical formulation is provided for the rotational, steady, inviscid, compressible motion in a solid rocket motor that is modeled as a slender porous chamber. The analysis overcomes some of the deficiencies encountered in previous work on the subject. The method that we employ consists of reducing the problem’s mass, momentum, energy, ideal gas, and isentropic relations into a single integral equation that can be solved numerically. Furthermore, Saint-Robert’s power law is used to link the pressure to the sidewall mass injection rate. At the outset, results are presented for the axisymmetric and planar porous chambers and compared to two closed-form analytical solutions developed under one-dimensional and two-dimensional, isentropic flow conditions, in addition to experimental data. The comparison is carried out assuming either uniformly distributed mass flux or constant injection speed along the porous wall. Our amended formulation is shown to agree with the one-dimensional solution obtained for the case of uniform wall mass flux and with the asymptotic approximation for the constant wall injection speed.

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.