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.

During the cure of polymer matrix composites, induced stresses develop due to shrinkage of the matrix material. Consequences of this can lead to shifting of the reinforcement, adversely affecting final properties of the material, or the induced stresses can alter the final geometry of the part. With the use of a new closed loop feedback program developed, residual stresses built up during cure were minimized. Experiments were performed using the EPON 828 resin with two types of reinforcement, carbon and glass fiber. The residual stress built up during the optimized cure cycle was compared with that produced during the lPanufacturer recommended 2-step cure cycle and isothermal cure cycles. Results for both fibers show a large reduction in stresses endured during cure for the optimized cure compared to typical stresses seen under isothermal and standard cure cycles. Static and dynamic testing were done on specimens and showed that the modulus and the glass transition temperatures of cured specimens were not significantly affected by the optimized cure cycles. Results also show that optimized cure cycles were of shorter duration compared to the standard cure cycles.

Thesis (M.App.Sc.) - Carleton University, 2009. / Includes bibliographical references (p.148-151). Also available in electronic format on the Internet.

<p> Crack propagation in a solid rocket motor environment is difficult to measure directly. This experimental and analytical study evaluated the viability of real-time radiography for detecting bore regression and propellant crack propagation speed. The scope included the quantitative interpretation of crack tip velocity from simulated radiographic images of a burning, center-perforated grain and actual real-time radiographs taken on a rapid-prototyped model that dynamically produced the surface movements modeled in the simulation. The simplified motor simulation portrayed a bore crack that propagated radially at a speed that was 10 times the burning rate of the bore. Comparing the experimental image interpretation with the calibrated surface inputs, measurement accuracies were quantified. The average measurements of the bore radius were within 3% of the calibrated values with a maximum error of 7%. The crack tip speed could be characterized with image processing algorithms, but not with the dynamic calibration data. The laboratory data revealed that noise in the transmitted X-Ray intensity makes sensing the crack tip propagation using changes in the centerline transmitted intensity level impractical using the algorithms employed.</p>

<p> The most important and crucial design feature while designing an Aircraft Electric Power Distribution System (EPDS) is reliability. In EPDS, the distribution of power is from top level generators to bottom level loads through various sensors, actuators and rectifiers with the help of AC &amp; DC buses and control switches. As the demands of the consumer is never ending and the safety is utmost important, there is an increase in loads and as a result increase in power management. Therefore, the design of an EPDS should be optimized to have maximum efficiency. This thesis discusses an integrated tool that is based on a Need Based Design method and Fault Tree Analysis (FTA) to achieve the optimum design of an EPDS to provide maximum reliability in terms of continuous connectivity, power management and minimum cost. If an EPDS is formulated as an optimization problem then it can be solved with the help of connectivity, cost and power constraints by using a linear solver to get the desired output of maximum reliability at minimum cost. Furthermore, the thesis also discusses the viability and implementation of the resulted topology on typical large aircraft specifications.</p>

<p> This study presents and investigates a novel hybrid fuel grain that reacts pyrophorically with gaseous oxidizer to achieve restart of a hybrid rocket motor propulsion system while reducing cost and handling concerns. This reactive fuel grain (RFG) relies on the pyrophoric nature of finely divided metal particles dispersed in a solid dicyclopentadiene (DCPD) binder, which has been shown to encapsulate air-sensitive additives until they are exposed to combustion gases. An RFG is thus effectively inert in open air in the absence of an ignition source, though the particles encapsulated within remain pyrophoric. In practice, this means that an RFG that is ignited in the vacuum of space and then extinguished will expose unoxidized pyrophoric particles, which can be used to generate sufficient heat to relight the propellant when oxidizer is flowed.</p><p> The experiments outlined in this work aim to develop a suitable pyrophoric material for use in an RFG, demonstrate pyrophoric relight, and characterize performance under conditions relevant to a hybrid rocket thruster. Magnesium, lithium, calcium, and an alloy of titanium, chromium, and manganese (TiCrMn) were investigated to determine suitability of pure metals as RFG additives. Additionally, aluminum hydride (AlH₃), lithium aluminum hydride (LiAlH₄), lithium borohydride (LiBH₄), and magnesium hydride (MgH₂) were investigated to determine suitability of metals hydrides as RFG additives or as precursors for pure-metal RFG additives. Pyrophoric metals have been previously investigated as additives for increasing the regression rate of hybrid fuels, but to the author&rsquo;s knowledge, these materials have not been specifically investigated for their ability to ignite a propellant pyrophorically.</p><p> Commercial research-grade metals were obtained as coarse powders, then ball-milled to attempt to reduce particle size below a critical diameter needed for pyrophoricity. Magnesium hydride was ball-milled and then cycled in a hydride cycling apparatus to attempt to fracture the particles through hydrogen sorption and thermal stresses. These powders were then tested for pyrophoricity with atmospheric and pure concentrations of oxygen. The TiCrMn powder was chosen as the material for evaluation of propellant performance, and was mixed with DCPD in various weight ratios to determine the required additive loading needed for pyrophoricity of the bulk propellant. Weight percentages of 10, 20, 30, and 50 wt.% TiCrMn were used to evaluate relight capability and propellant performance, and weight loadings of 50, 70, and 90 wt.% TiCrMn were used to evaluate approximate maximum loading possible without rendering the propellant structurally unsound. Propellant tests were conducted in an opposed flow burner apparatus for sub-scale regression rate and relight experiments, and an optically accessible cylindrical combustion chamber (OCC) that allows high speed cameras to record the regressing propellant surface during combustion. Gaseous oxygen (GOX) was used as an oxidizer for all tests due to its ready availability and common use as a hybrid rocket oxidizer. Opposed flow burner experiments are an inexpensive means of rapidly testing various propellant formulations at different conditions, whereas OCC tests are useful for obtaining realistic data on how an RFG would likely operate as part of a propulsion system.</p><p> Relight in the opposed flow burner was attempted by cycling oxygen and nitrogen flows with carefully timed solenoid valves to initiate and extinguish combustion, and to control the slow diffusion of oxygen to the surface of the propellant, which would render the TiCrMn non-pyrophoric. The opposed flow burner experiments did not conclusively demonstrate the pyrophoric relight capability of the RFG propellant due in part to the persistence of hot spots between oxygen and purge nitrogen cycles, as determined by high-speed imaging in the near infrared range. An opposed flow burner apparatus was then constructed within a vacuum chamber assembly thus preventing atmospheric oxygen from diffusing to the propellant surface, but these tests did not demonstrate pyrophoric relight. Future work is proposed to evaluate the effect of pyrophoric particle size in order to determine the role ignition delay of each particle has in the relight capability of RFGs.</p><p> OCC experiments were conducted at a low and high GOX mass flux of approximately 150 and 300 kg/s/m², respectively, at a nominal chamber pressure of 150 psia. Four strand compositions were used: pure DCPD, 30 wt.% pyrophoric TiCrMn powder with average particle diameters of approximately 1-10 microns, 30 wt.% oxidized TiCrMn powder with average particle diameters of approximately 1-10 microns, and 30 wt.% TiCrMn powder with average particle diameters of approximately 1-4 mm. Regression rate was measure by weight loss, average web thickness change at three axial locations on the strand, and through time-resolved tracking of the regressing propellant surface via high speed video. While visual observations suggest that the addition of TiCrMn significantly increases regression rate, initial data do not show a significant trend. Additionally, it is observed that the oxidized TiCrMn strands regress at the same rate as those loaded with pyrophoric TiCrMn, suggesting that erosive burning and heat addition of the added metal may be the cause of the observed increase in regression rate. The data are too sparse to make conclusions about the effect of particle size on regression rate, so further tests are recommended to develop a significant data set for the effect of pyrophoricity and particle size on regression rate. The test article was damaged at the end of the regression rate experimental campaign, which precluded the collection of relight data that was planned for strands loaded with 50 wt.% TiCrMn particles with an average diameter of approximately 1-4 mm. Though further tests are needed to demonstrate pyrophoric relight of an RFG, the current work establishes a baseline for RFG performance and suggests that pyrophoric relight is possible by tailoring the particle size of the pyrophoric metal additive to control heat release and ignition delay.</p>

The streamwise growth rate of the mixing layer downstream of a splitter plate with a 60° swept lambda notch and velocity ratio of 0.5 was spanwise independent and self-preserving once the velocity deficit from the splitter plate had disappeared. It was hypothesized the baseline mixing layer had a rapidly divergent growth component normal, and a non-divergent component parallel to the trailing edge of the notch. The streamwise growth rate of flow was not dependent on the spanwise dependence of the boundary layer. For forced cases the interaction of waves propagating from the oscillating flaperons influenced the streamwise growth rate of mixing layer. The forced case had a divergent growth component parallel to the trailing edge associated with streamwise rib vortices. The initial growth stages of the forced cases of this turbulent 3D mixing layer were dependent on the local Strouhal Number; similar to certain 2D cases.