Rotating detonation combustors and their similarities to rocket instabilities

A quadrotor UAV including ruggedized, integral-battery, load-bearing body, two arms on the load-bearing body, each arm having two rotors, a control module mounted on the load-bearing body, a payload module mounted on the control module, and skids configured as landing gear. The two arms are replaceable with arms having wheels for ground vehicle use, with arms having floats and props for water-surface use, and with arms having pitch-controlled props for underwater use. The control module is configured to operate as an unmanned aerial vehicle, an unmanned ground vehicle, an unmanned (water) surface vehicle, and an unmanned underwater vehicle, depending on the type of arms that are attached.

Reconfigurable battery-operated vehicle system

Recent advances in new oxidizers for solid rocket propulsion

Inearlierworkithasbeenshownthatspheresofvarioussizescanberandomlypackedtosimulatethemorphology ofammoniumperchlorate (AP)/binderheterogeneouspropellants.Amodelisnowformulated in whichpropellants dee nedinthiswaycanbeburnt,allowingforcompletecouplingbetweenthegas-phasephysics,thecondensed-phase physics, and the unsteady nonuniform regression of the propellant surface. The gas-phase kinetics is represented by a two-step model with parameters e tted to experimental data for the one-dimensional combustion of pure AP and the one-dimensional combustion of a e ne-AP/binder blend. In the discussion of the pure AP e ts, the issue of intrinsic instabilities arises, and these are explored using nonlinear direct numerical simulation and a numerical linear-stability strategy. Results for two-dimensional heterogeneous burning show that surface regions where the AP and binder can mix regress more rapidly than those dominated by AP; local extinction can occur where the binder dominates. Time variations in the integrated e ux-based equivalence ratio are shown to be large. Variations in the average burning rate with pressure are consistent with the (three-dimensional ) experimental record. Nomenclature cp = specie c heat D1;2 = reaction rate constants dj = diameter of the jth particle class E = activation energy f = surface function M = mass e ux Nj = number in the jth particle class n1;2 = pressure exponents in the reaction rates

Heterogeneous Propellant Combustion

In this article we point to the ways in which engineering and communication disciplines can work together to ensure the ABET criterion that encompasses effective communication is represented in engineering curricula. Drawing upon examples from several universities in North America, we offer useful portraits of writing across the curriculum approaches, interdisciplinary courses, integrated programs, and a variety of support systems including writing and communication centers and online resources. As we develop increased awareness of the importance of including communication instruction in engineering curricula, the variety of possibilities presented in this article can help us integrate communication and engineering education.

Integrating Communication and Engineering Education: A Look at Curricula, Courses, and Support Systems

The low regression rates of classic hybrid rocket fuels lead to large internal ports that limit potential applications. This experimental study investigated the increase in regression rate that results from adding a solid oxidizer and a catalyst to a hybrid fuel grain. The configuration is named a "mixed hybrid" hybrid to signify solid oxidizer and catalyst in the grain. A design of experiments approach guided fuel formulation to systematically control levels of ammonium perehlorate from 25% to 30%, ferric oxide from 0 to 5%, and hydroxyl-terminated polybutadiene from 70% to 75%. The 1.5-in. diam. port, 12-in. long center perforated grains were burned with gaseous oxygen at pressure levels from 150 to 550 psig and port flux levels from 0.1 to 0.4 lbm/s-in. 2 . The results show that the mixed hybrid propellants burn as a function of both pressure and mass flux. A grain formulation having 27.5 % ammonium perchlorate and 2.5% ferric oxide provided the maximum burning rate augmentation (447%) among the formulations tested.

Regression rates study of mixed hybrid propellants

The UAH Propulsion Research Center (PRC) is in its 29 year at the University of Alabama in Huntsville (UAH). The mission of the Propulsion Research Center is to provide an environment that connects the academic research community with the needs and concerns of the propulsion community while promoting an interdisciplinary approach to solving propulsion problems. This paper summarizes recent metrics from academic and research programs that are associated with the PRC. The emphasis this year is describing the graduate student production supported by the PRC over the past 29 years. This assessment identified over 270 theses and dissertations into 9 different propulsion categories. The results accumulated into 207 master’s and 67 Ph.D.s with the highest production rate being 21 degrees per year in 2010. The documents are distributed into categories representing conventional propulsion, advanced propulsion, missile design, airbreathing propulsion, Propulsion Testing, and Propulsion Systems Engineering. About 83% of the graduates are estimated to be U.S. citizens and there was a 20% female graduation rate. For the current year the total research expenditures from fifteen different agencies are anticipated to rise to $3.3 million which is a 33% compared to last year. PRC researchers published over 130 papers last year. The UAH student launch team placed 3 in the NASA Student Launch national competition this year. The PRC continues to be a resource to accomplish high-quality research, education, and workforce development in propulsion and energy.

Correction: Propulsion Research and Academic Programs at the University of Alabama in Huntsville - PRC Graduate Student Production History

This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

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A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

Laboratory-scale hybrid rocket motors often provide the means to evaluate the burning rates of new fuel formulations. While the issue of scaling to larger sizes can be significant, the uncertainty level of the laboratory-scale results also needs evaluation. This work quantitatively evaluates the uncertainty of key experimental results for a particular hybrid rocket motor. The scope includes calculation of the uncertainties in the fuel regression rate, oxidizer flux, motor characteristic velocity, and the oxidizer-to-fuel mass ratio for a laboratory-sca le motor. The motor burned gaseous oxygen and hydroxyl-terminated polybutadiene at conditions characteristic of a preburner. The accepted uncertainty methodology of Coleman and Steele established the approach for the study. The results show that the typical uncertainty values are ±8.7% hi the determination of the fuel regression rate, ±10.8% in the oxidizer flux, ±10.5% for motor characteristic velocity, and ±9.2% for the oxidizer-to-fuel ratio. Measurement uncertainties hi the diameters of the initial fuel grain, initial port, and nozzle throat contributed significantly to these conclusions. Conceptual biases in determining the burn duration, the average chamber pressure, and the average burning rate have a significant influence. This evaluation revealed the important measurement limitations and potential improvements for the particular test conditions and data reduction equations used. However, the methodology presented is in generally applicable to hybrid rocket testing.

Laboratory-scale hybrid rocket motor uncertainty analysis

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 axial-injection, 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 bas...

Robert A. Frederick

Papers

Regression rates study of mixed hybrid propellants

Correction: Propulsion Research and Academic Programs at the University of Alabama in Huntsville - PRC Graduate Student Production History

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

Laboratory-scale hybrid rocket motor uncertainty analysis

Testing and Modeling of a Porous Axial-Injection, End-Burning Hybrid Motor