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Dextre R.A.,Propulsion Research Center | Xu K.G.,University of Alabama in Huntsville
51st AIAA/SAE/ASEE Joint Propulsion Conference | Year: 2015

In this paper, a microstrip split ring resonator microwave-induced plasma source is developed. The goal of this work is to implement the resonator into a micropropulsion system. This project evaluated three designs of the resonator to understand how the width of the ring microstrip can impact the microplasma properties. Each resonator was designed and fabricated to operate at ~961 MHz. Simulations of the electric fields were performed. Single Langmuir probe measurements were done at 10 Torr to obtain the electron density, temperature, and plasma potential. The microwave power was varied from 10-15 W. The results show that the 1.5 mm wide resonator has the greatest electron temperature compared to the 0.5 mm and 1 mm. The electron density is also greatest for the 1.5 mm device. Ion density is higher for the 0.5 mm compared to the 1 mm and 1.5 mm. The simulations show that with an increase in ring width, there is an increase in electric field in the microstrip of the resonator. These results help to understand the relationship between microstrip width and the microplasma produced. As a result of this research, the proper split ring resonator source is determined for implementation into a microthruster. Overall, this research will lead to the production of an optimized microplasma source for microthruster applications. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Source

Hitt M.A.,University of Alabama in Huntsville | Hitt M.A.,State University of New York at Buffalo | Frederick R.A.,University of Alabama in Huntsville | Frederick R.A.,Propulsion Research Center
Journal of Propulsion and Power | Year: 2016

Testing was conducted using polyethylene as the porous fuel and gaseous oxygen as the oxidizer. Nominal test articles were tested using 100, 50, and 15 μm pore sizes. Pressures tested ranged from atmospheric to 1194 kPa, and oxidizer injection velocities ranged from 35 to 80 m/s. Regression rates were determined using pretest and posttest length measurements of the solid fuel. Experimental results demonstrated that the regression rate of the porous axialinjection, end-burning hybrid was a function of the chamber pressure, as opposed to the oxidizer mass flux typical in conventional hybrids. Regression rates ranged from approximately 0.65 mm/s at atmospheric pressure to 7.74 mm/s at 1194 kPa. The analytical model was developed based on a standard ablative model modified to include oxidizer flow through the grain. The heat transfer from the flame was primarily modeled using an empirically determined flame coefficient that included all heat transfer mechanisms in one term. An exploratory flame model based on the granular diffusion flame model used for solid rocket motors was also adapted for comparison with the empirical flame coefficient. This model showed agreement with the experimental results, indicating that it has potential for giving insight into the flame structure in this motor configuration. Copyright ©2015 by M. A. Hitt and R. A. Frederick Jr. Source

Frederick R.A.,Propulsion Research Center
50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference 2014 | Year: 2014

Advances in rocket propulsion technology have historically relied on a healthy interaction among members of the university, industry, and government communities. The National Institute of Rocket Propulsion Systems commissioned a series of three Academic Workshops from 2011 to 2014 to gather information about the current status of research at universities and recommend strategies for successful collaborations. This paper summarizes the activities of these workshops. A snapshot of propulsion research at universities is summarized from Workshop I. Workshop II produced a series of health metrics by which one could monitor the well-being of university propulsion programs and a description of how individual universities organized their efforts. In Workshop III, representatives of the university, industry, and government examined three case studies on larger university-based activities, which interacted heavily with government and involved industry. Participants assessed desirable attributes such as continuity of support, sustainability, industry mentorship, multi-disciplinary collaboration, and shared curricula. University coalitions that operate under non-profit research and graduate education institute appear to be an excellent strategy for current and future sustainable university, government, and industry programs in rocket propulsion systems. © 2014 by Robert A. Frederick, Jr. Source

Butt A.,University of Alabama in Huntsville | Frederick R.A.,University of Alabama in Huntsville | Frederick R.A.,Propulsion Research Center | Denny M.,University of Alabama in Huntsville
51st AIAA/SAE/ASEE Joint Propulsion Conference | Year: 2015

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, centerperforated grain and actual real-time radiographs taken on a rapid-prototyped model that dynamically produces 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 and 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 agorithms employed. This paper details the design of rapid-protyped test models and the experimental x-ray test setup used, along with the simulation techniques employed. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Source

Bennewitz J.W.,University of Alabama in Huntsville | Bennewitz J.W.,Propulsion Research Center | Frederick R.A.,University of Alabama in Huntsville | Frederick R.A.,Propulsion Research Center | And 4 more authors.
Journal of Propulsion and Power | Year: 2015

This research investigation encompasses experimental tests demonstrating the control of a high-frequency combustion instability by acoustically modulating the propellant flow. This investigation complements an accompanying theoretical study implementing linear modal analysis [Bennewitz, J. W., Rani, S. L., Cranford, J. T., and Frederick, R. A., Jr., "Combustion Instability Control Through Acoustic Modulation at the Inlet Boundary: Analysis," Journal of Propulsion and Power, (to be published)].Amodel rocket combustor burned gaseous oxygen and methane using a single-element pentad-style injector. Flow conditions were established that spontaneously excited a 2430 Hz first longitudinal combustion oscillation at an amplitude up to p′/pc ∼ 4%. An acoustic speaker was placed at the base of the oxidizer supply line to modulate the flow and alter the oscillatory behavior of the combustor. Two speaker modulation approaches were investigated: 1) bands of white noise, and 2) pure sinusoidal tones. The first approach adjusted 500 Hz bands of white noise ranging from 0-500 to 2000-2500 Hz, whereas the second approach implemented individual harmonic signals with arbitrary phase swept from 500 to 2500 Hz. The results show that, above a modulation signal amplitude threshold, both approaches suppressed 95+%of the spontaneous combustion oscillation. By increasing the applied signal amplitude, a wider frequency range of instability suppression became present for these two acoustic modulation approaches. Thus, this work further supports the strategic application of acoustic modulation within an injector as a potential method to control high-frequency combustion instabilities for liquid rocket engine applications. © 2015 by JohnW. Bennewitz, Robert A. Frederick, Jr., Jacob T. Cranford, and David M. Lineberry. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Source

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