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HIGH PRESSURE TESTING OF COMPOSITE SOLID ROCKET PROPELLANT MIXTURES: BURNER FACILITY CHARACTERIZATION
Title
HIGH PRESSURE TESTING OF COMPOSITE SOLID ROCKET PROPELLANT MIXTURES: BURNER FACILITY CHARACTERIZATION.
Creator
Carro, Rodolphe, Petersen, Eric, University of Central Florida
Abstract / Description
Much Research on composite solid propellants has been performed over the past few decades and much progress has been made, yet many of the fundamental processes are still unknown, and the development of new propellants remains highly empirical. Ways to enhance the performance of solid propellants for rocket and other applications continue to be explored experimentally, including the effects of various additives and the impact of fuel and oxidizer particle sizes on burning behavior. One...
Show moreMuch Research on composite solid propellants has been performed over the past few decades and much progress has been made, yet many of the fundamental processes are still unknown, and the development of new propellants remains highly empirical. Ways to enhance the performance of solid propellants for rocket and other applications continue to be explored experimentally, including the effects of various additives and the impact of fuel and oxidizer particle sizes on burning behavior. One established method to measure the burning rate of composite propellant mixtures in a controlled laboratory setting is to use a constant-volume pressure vessel, or strand burner. To provide high-pressure burn rate data at pressures up to 360 atm, the authors have installed, characterized and improved a strand burner facility at the University of Central Florida. Details on the facility and its improvements, the measurement procedures, and the data reduction and interpretation are presented. Two common HTPB/AP propellant mixtures were tested in the original strand burner. The resulting burn rates were compared to data from the literature with good agreement, thus validating the facility and related test techniques, the data acquisition, data reduction and interpretation. After more than 380 successful recordings, an upgraded version of the strand burner, was added to the facility. The details of Strand Burner II, its improvements over Strand Burner I, and its characterization study are presented.
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Date Issued
2007
Identifier
CFE0001979, ucf:47427
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0001979
SOUNDING ROCKET REDESIGN AND OPTIMIZATION FOR PAYLOAD EXPANSION AND IN FLIGHT TELEMETRY TRANSMITTAL
Title
SOUNDING ROCKET REDESIGN AND OPTIMIZATION FOR PAYLOAD EXPANSION AND IN FLIGHT TELEMETRY TRANSMITTAL.
Creator
Huffman, Matthew, Chew, Larry, University of Central Florida
Abstract / Description
Due to renewed interest in the sub orbital rocket program of the Florida Space Authority and a surplus of Super Loki Sounding Rockets, an effort to improve the usefulness of this surplus is herein undertaken. Currently, the capacity of the payload compartment in the upper stage of the Super Loki system is very limited. A redesign of the upper stage will allow larger and more versatile payloads to be carried, assuming the appropriate design parameters are met. It has therefore been undertaken...
Show moreDue to renewed interest in the sub orbital rocket program of the Florida Space Authority and a surplus of Super Loki Sounding Rockets, an effort to improve the usefulness of this surplus is herein undertaken. Currently, the capacity of the payload compartment in the upper stage of the Super Loki system is very limited. A redesign of the upper stage will allow larger and more versatile payloads to be carried, assuming the appropriate design parameters are met. It has therefore been undertaken to create a design procedure that is comprehensive in scope in order to affect this redesign. This procedure includes five major components. These are the separation of the upper and lower stages, the stability of the vehicle, the altitude and velocity of the rocket, the mechanical loading and finally the aerodynamic heating. Semi-empirical methods were used whenever possible to allow comparison with experimental data. This procedure revealed that larger diameter upper stages might be used up to a reasonable maximum of four inches. The four-inch modification is found to be stable as were the smaller modifications considered. The altitude and velocity of the rocket were found via a simple Eulerian time stepping scheme resulting in an estimate of approximately 148,000ft for the four-inch dart. The mechanical loading analysis allowed for the material selection for the rocket components. Reinforced steel fins and carbon fiber tubing, for the payload section, are adequate to meet expected mechanical loads, those being, 16000psi for the fin section due to launcher forces, 22800psi for compressive plus torsion forces on the composite section and 18000psi for the ejection stresses. An ablative coating is considered necessary to counteract the 760ºF temperatures along the composite tube.
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Date Issued
2005
Identifier
CFE0000546, ucf:46440
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0000546
DESIGN AND FABRICATION OF A FULL-FEATURED LABSCALE HYBRID ROCKET ENGINE
Title
DESIGN AND FABRICATION OF A FULL-FEATURED LABSCALE HYBRID ROCKET ENGINE.
Creator
Platt, Kyle, Petersen, Eric, University of Central Florida
Abstract / Description
The design, development, integration and testing of a full-featured, Lab-Scale Hybrid Rocket Engine was not only envisioned to be the chosen method of putting student payloads into space, but to be an invaluable teaching resource. The subject of the present thesis is the analysis, design, development, integration and demonstration of a lab-scale hybrid rocket motor. The overarching goal of this project was to establish a working developmental lab model from which further research can be...
Show moreThe design, development, integration and testing of a full-featured, Lab-Scale Hybrid Rocket Engine was not only envisioned to be the chosen method of putting student payloads into space, but to be an invaluable teaching resource. The subject of the present thesis is the analysis, design, development, integration and demonstration of a lab-scale hybrid rocket motor. The overarching goal of this project was to establish a working developmental lab model from which further research can be accomplished. The lab model was specifically designed to use a fuel source that could be studied in normal laboratory conditions. As such, the rocket engine was designed to use Hydroxyl Terminated Polybutadiene as the fuel and Liquid Nitrous Oxide as the oxidizer. Developing the rocket engine required the usage of several electronics modules and a software package. The custom-designed electronics modules were a Signal Conditioning & Data Amplification Interface and a Data Acquisition Network. The software package was coded in Visual Basic (VB). A MathCAD regression rate computer model was designed and written to geometrically constrain the engine design. Further, the computer model allowed for the "what-if" situations to be evaluated. Using ProPep, solutions to the Equilibrium Thermodynamics Equations for the fuel and oxidizer mixture were obtained. The resultants were used as initial input to the computer model for predicting the lab-scale rocket's Chamber Pressure, Chamber Temperature, Ratio of Specific Heats and Molecular Weight. Details on the model, the rocket hardware, and the successful test firing are provided.
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Date Issued
2006
Identifier
CFE0000972, ucf:46714
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0000972
THE RISE OF ASIA'S DEMOCRATIC SPACE POWERS: HOW JAPAN AND INDIA BECAME THE NEXT SPACE POWERS IN THE TWENTY-FIRST CENTURY.
Title
THE RISE OF ASIA'S DEMOCRATIC SPACE POWERS: HOW JAPAN AND INDIA BECAME THE NEXT SPACE POWERS IN THE TWENTY-FIRST CENTURY.
Creator
Kunze, Shane, Handberg, Roger, University of Central Florida
Abstract / Description
Since the end of World War II the world has seen several nations expand into the space age. Also after the Second World War, the Cold War began and many nations found themselves allying themselves with either the hegemony of the West or the Communists. Space was no exception in this dilemma, as weaker nations began to develop their own indigenous space programs and had technological diffusion from one of the hegemonies. Japan and India are two democracies that both sought support for their...
Show moreSince the end of World War II the world has seen several nations expand into the space age. Also after the Second World War, the Cold War began and many nations found themselves allying themselves with either the hegemony of the West or the Communists. Space was no exception in this dilemma, as weaker nations began to develop their own indigenous space programs and had technological diffusion from one of the hegemonies. Japan and India are two democracies that both sought support for their indigenous space programs from the west, particularly from the U.S. These two nations emerged from poverty and a broken infrastructure during the 1950s and have grown over the last sixty years into two of the most advanced space-faring nations in the world. These two nations have overcome several external and internal factors ranging from Communist expansion to bureaucratic strife. Japan and India have been and remain the two leading democratic nations in Asia that have risen to the rank of space power.
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Date Issued
2012
Identifier
CFH0004170, ucf:44858
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFH0004170
Investigation into the Feasibility of Adding Turbulators to Rocket Combustion Chamber Cooling Channels Using a Conjugate Heat Transfer Analysis
Title
Investigation into the Feasibility of Adding Turbulators to Rocket Combustion Chamber Cooling Channels Using a Conjugate Heat Transfer Analysis.
Creator
Buchanan, Tyler, Kapat, Jayanta, Raghavan, Seetha, Ghosh, Ranajay, University of Central Florida
Abstract / Description
A conjugate heat transfer analysis will be carried out to simulate an 89 kN thrust chamber hydrogen cooling channel, to determine the feasibility of adding turbulators to the combustion chamber cooling channels at various parameters such as angle, pitch, and height of the turbulator. An existing regeneratively cooled chamber environment is simulated and used as a baseline case to be compared against. The new design includes using ribbed turbulators or delta wedges in the cooling channels to...
Show moreA conjugate heat transfer analysis will be carried out to simulate an 89 kN thrust chamber hydrogen cooling channel, to determine the feasibility of adding turbulators to the combustion chamber cooling channels at various parameters such as angle, pitch, and height of the turbulator. An existing regeneratively cooled chamber environment is simulated and used as a baseline case to be compared against. The new design includes using ribbed turbulators or delta wedges in the cooling channels to increase the heat transfer on the channel hot wall (wall adjacent to the hot gas wall) and on the two channel sidewalls. With a higher heat transfer coefficient, the sidewalls behave like fins for heat transfer and participate more in the overall heat transfer process in the channel. Efficient rib and wedge geometries are chosen based on previous investigations. A conjugate heat transfer analysis is performed using a straight duct with the rib and wedge geometries included, with boundary conditions similar to those found in the combustion chamber, to provide thermal hydraulic performance data at numerous turbulator configurations. The baseline channel's maximum hot wall temperature is the target maximum hot wall temperature that is desired to be reduced. The goal is to reduce the hot gas side wall temperature at a minimal cost in pressure drop.
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Date Issued
2018
Identifier
CFE0007160, ucf:52320
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0007160
DESIGN AND IMPLEMENTATION OF AN EMISSION SPECTROSCOPY DIAGNOSTIC IN A HIGH-PRESSURE STRAND BURNER FOR THE STUDY OF SOLID PROPELLANT COMBUSTION
Title
DESIGN AND IMPLEMENTATION OF AN EMISSION SPECTROSCOPY DIAGNOSTIC IN A HIGH-PRESSURE STRAND BURNER FOR THE STUDY OF SOLID PROPELLANT COMBUSTION.
Creator
Arvanetes, Jason, Petersen, Eric, University of Central Florida
Abstract / Description
The application of emission spectroscopy to monitor combustion products of solid rocket propellant combustion can potentially yield valuable data about reactions occurring within the volatile environment of a strand burner. This information can be applied in the solid rocket propellant industry. The current study details the implementation of a compact spectrometer and fiber optic cable to investigate the visible emission generated from three variations of solid propellants. The grating was...
Show moreThe application of emission spectroscopy to monitor combustion products of solid rocket propellant combustion can potentially yield valuable data about reactions occurring within the volatile environment of a strand burner. This information can be applied in the solid rocket propellant industry. The current study details the implementation of a compact spectrometer and fiber optic cable to investigate the visible emission generated from three variations of solid propellants. The grating was blazed for a wavelength range from 200 to 800 nm, and the spectrometer system provides time resolutions on the order of 1 millisecond. One propellant formula contained a fine aluminum powder, acting as a fuel, mixed with ammonium perchlorate (AP), an oxidizer. The powders were held together with Hydroxyl-Terminated-Polybutadiene (HTPB), a hydrocarbon polymer that is solidified using a curative after all components are homogeneously mixed. The other two propellants did not contain aluminum, but rather relied on the HTPB as a fuel source. The propellants without aluminum differed in that one contained a bimodal mix of AP. Utilizing smaller particle sizes within solid propellants yields greater surface area contact between oxidizer and fuel, which ultimately promotes faster burning. Each propellant was combusted in a controlled, non-reactive environment at a range of pressures between 250 and 2000 psi. The data allow for accurate burning rate calculations as well as an opportunity to analyze the combustion region through the emission spectroscopy diagnostic. It is shown that the new diagnostic identifies the differences between the aluminized and non-aluminized propellants through the appearance of aluminum oxide emission bands. Anomalies during a burn are also verified through the optical emission spectral data collected.
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Date Issued
2006
Identifier
CFE0000971, ucf:46694
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0000971
REDUCTION OF VORTEX-DRIVEN OSCILLATIONS IN A SOLID ROCKET MOTOR COLD FLOW SIMULATION THROUGH ACTIVE CONTROL
Title
REDUCTION OF VORTEX-DRIVEN OSCILLATIONS IN A SOLID ROCKET MOTOR COLD FLOW SIMULATION THROUGH ACTIVE CONTROL.
Creator
Ward, Jami, Leonessa, Alexander, University of Central Florida
Abstract / Description
Control of vortex-driven instabilities was demonstrated via a scaled-down, cold-flow simulation that modeled closed-end acoustics. When vortex shedding frequencies couple with the natural acoustic modes of a choked chamber, potentially damaging low-frequency instabilities may arise. Although passive solutions can be effective, an active control solution is preferable. An experiment was performed to demonstrate an active control scheme for the reduction of vortex-driven oscillations. A non...
Show moreControl of vortex-driven instabilities was demonstrated via a scaled-down, cold-flow simulation that modeled closed-end acoustics. When vortex shedding frequencies couple with the natural acoustic modes of a choked chamber, potentially damaging low-frequency instabilities may arise. Although passive solutions can be effective, an active control solution is preferable. An experiment was performed to demonstrate an active control scheme for the reduction of vortex-driven oscillations. A non-reacting experiment using a primary flow of air, where both the duct exit and inlet are choked, simulated the closed-end acoustics. Two plates, separated by 1.27 cm, produced the vortex shedding phenomenon at the chamber's first longitudinal mode. Two active control schemes, closed-loop and open-loop, were studied via a cold-flow simulation for validating the effects of reducing vortex shedding instabilities in the system. Actuation for both control schemes was produced by using a secondary injection method. The actuation system consisted of pulsing compressed air from a modifed, 2-stroke model airplane engine, controlled and powered by a DC motor. The use of open-loop only active control was not highly effective in reducing the amplitude of the first longitudinal acoustic mode, near 93 Hz, when the secondary injection was pulsed at the same modal frequency. This was due to the uncontrolled phasing of the secondary injection system. A Pulse Width Modulated (PWM) signal was added to the open-loop control scheme to correct for improper phasing of the secondary injection flow relative to the primary flow. This addition allowed the motor speed to be intermittently increased to a higher RPM before returning to the desired open-loop control state. This proved to be effective in reducing the pressure disturbance by approximately 46%. A closed-loop control scheme was then test for its effectiveness in controlling the phase of the secondary injection. Feedback of the system's state was determined by placing a dynamic pressure transducer near the chamber exit. Closed-loop active control, using the designed secondary injection system, was proven as an effective means of reducing the problematic instabilities. A 50% reduction in the FFT RMS amplitude was realized by utilizing a Proportional-Derivative controller to modify the phase of the secondary injection.
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Date Issued
2006
Identifier
CFE0000920, ucf:46728
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0000920
CONJUGATE HEAT TRANSFER ANALYSIS OF COMBINED REGENERATIVE AND DISCRETE FILM COOLING IN A ROCKET NOZZLE
Title
CONJUGATE HEAT TRANSFER ANALYSIS OF COMBINED REGENERATIVE AND DISCRETE FILM COOLING IN A ROCKET NOZZLE.
Creator
Pearce, Charlotte M, Kapat, Jayanta, University of Central Florida
Abstract / Description
Conjugate heat transfer analysis has been carried out on an 89kN thrust chamber in order to evaluate whether combined discrete film cooling and regenerative cooling in a rocket nozzle is feasible. Several cooling configurations were tested against a baseline design of regenerative cooling only. New designs include combined cooling channels with one row of discrete film cooling holes near the throat of the nozzle, and turbulated cooling channels combined with a row of discrete film cooling...
Show moreConjugate heat transfer analysis has been carried out on an 89kN thrust chamber in order to evaluate whether combined discrete film cooling and regenerative cooling in a rocket nozzle is feasible. Several cooling configurations were tested against a baseline design of regenerative cooling only. New designs include combined cooling channels with one row of discrete film cooling holes near the throat of the nozzle, and turbulated cooling channels combined with a row of discrete film cooling holes. Blowing ratio and channel mass flow rate were both varied for each design. The effectiveness of each configuration was measured via the maximum hot gas-side nozzle wall temperature, which can be correlated to number of cycles to failure. A target maximum temperature of 613K was chosen. Combined film and regenerative cooling, when compared to the baseline regenerative cooling, reduced the hot gas side wall temperature from 667K to 638K. After adding turbulators to the cooling channels, combined film and regenerative cooling reduced the temperature to 592K. Analysis shows that combined regenerative and film cooling is feasible with significant consequences, however further improvements are possible with the use of turbulators in the regenerative cooling channels.
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Date Issued
2016
Identifier
CFH2000138, ucf:45923
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFH2000138
DESIGN OPTIMIZATION OF SOLID ROCKET MOTOR GRAINS FOR INTERNAL BALLISTIC PERFORMANCE
Title
DESIGN OPTIMIZATION OF SOLID ROCKET MOTOR GRAINS FOR INTERNAL BALLISTIC PERFORMANCE.
Creator
Hainline, Roger, Nayfeh, Jamal, University of Central Florida
Abstract / Description
The work presented in this thesis deals with the application of optimization tools to the design of solid rocket motor grains per internal ballistic requirements. Research concentrated on the development of an optimization strategy capable of efficiently and consistently optimizing virtually an unlimited range of radial burning solid rocket motor grain geometries. Optimization tools were applied to the design process of solid rocket motor grains through an optimization framework developed to...
Show moreThe work presented in this thesis deals with the application of optimization tools to the design of solid rocket motor grains per internal ballistic requirements. Research concentrated on the development of an optimization strategy capable of efficiently and consistently optimizing virtually an unlimited range of radial burning solid rocket motor grain geometries. Optimization tools were applied to the design process of solid rocket motor grains through an optimization framework developed to interface optimization tools with the solid rocket motor design system. This was done within a programming architecture common to the grain design system, AML. This commonality in conjunction with the object-oriented dependency-tracking features of this programming architecture were used to reduce the computational time of the design optimization process. The optimization strategy developed for optimizing solid rocket motor grain geometries was called the internal ballistic optimization strategy. This strategy consists of a three stage optimization process; approximation, global optimization, and highfidelity optimization, and optimization methodologies employed include DOE, genetic algorithms, and the BFGS first-order gradient-based algorithm. This strategy was successfully applied to the design of three solid rocket motor grains of varying complexity. The contributions of this work was the development and application of an optimization strategy to the design process of solid rocket motor grains per internal ballistic requirements.
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Date Issued
2006
Identifier
CFE0001236, ucf:46929
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0001236
NUMERICAL COMPUTATIONS FOR PDE MODELS OF ROCKET EXHAUST FLOW IN SOIL
Title
NUMERICAL COMPUTATIONS FOR PDE MODELS OF ROCKET EXHAUST FLOW IN SOIL.
Creator
Brennan, Brian, Moore, Brian, University of Central Florida
Abstract / Description
We study numerical methods for solving the nonlinear porous medium and Navier-Lame problems. When coupled together, these equations model the flow of exhaust through a porous medium, soil, and the effects that the pressure has on the soil in terms of spatial displacement. For the porous medium equation we use the Crank-Nicolson time stepping method with a spectral discretization in space. Since the Navier-Lame equation is a boundary value problem, it is solved using a finite element method...
Show moreWe study numerical methods for solving the nonlinear porous medium and Navier-Lame problems. When coupled together, these equations model the flow of exhaust through a porous medium, soil, and the effects that the pressure has on the soil in terms of spatial displacement. For the porous medium equation we use the Crank-Nicolson time stepping method with a spectral discretization in space. Since the Navier-Lame equation is a boundary value problem, it is solved using a finite element method where the spatial domain is represented by a triangulation of discrete points. The two problems are coupled by using approximations of solutions to the porous medium equation to define the forcing term in the Navier-Lame equation. The spatial displacement solutions can be used to approximate the strain and stress imposed on the soil. An analysis of these physical properties shows whether or not the material ceases to act as an elastic material and instead behaves like a plastic which will tell us if the soil has failed and a crater has formed. Analytical as well as experimental tests are used to find a good balance for solving the porous medium and Navier-Lame equations both accurately and efficiently.
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Date Issued
2010
Identifier
CFE0003217, ucf:48565
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0003217
Factors Affecting Systems Engineering Rigor in Launch Vehicle Organizations
Title
Factors Affecting Systems Engineering Rigor in Launch Vehicle Organizations.
Creator
Gibson, Denton, Karwowski, Waldemar, Rabelo, Luis, Kotnour, Timothy, Kern, David, University of Central Florida
Abstract / Description
Systems engineering is a methodical multi-disciplinary approach to design, build, and operate complex systems. Launch vehicles are considered by many extremely complex systems that have greatly impacted where the systems engineering industry is today. Launch vehicles are used to transport payloads from the ground to a location in space. Satellites launched by launch vehicles can range from commercial communications to national security payloads. Satellite costs can range from a few million...
Show moreSystems engineering is a methodical multi-disciplinary approach to design, build, and operate complex systems. Launch vehicles are considered by many extremely complex systems that have greatly impacted where the systems engineering industry is today. Launch vehicles are used to transport payloads from the ground to a location in space. Satellites launched by launch vehicles can range from commercial communications to national security payloads. Satellite costs can range from a few million dollars to billions of dollars. Prior research suggests that lack of systems engineering rigor as one of the leading contributors to launch vehicle failures. A launch vehicle failure could have economic, societal, scientific, and national security impacts. This is why it is critical to understand the factors that affect systems engineering rigor in U.S. launch vehicle organizations.The current research examined organizational factors that influence systems engineering rigor in launch vehicle organizations. This study examined the effects of the factors of systems engineering culture and systems engineering support on systems engineering rigor. Particularly, the effects of top management support, organizational commitment, systems engineering support, and value of systems engineering were examined. This research study also analyzed the mediating role of systems engineering support between top management support and systems engineering rigor, as well as between organizational commitment and systems engineering rigor. A quantitative approach was used for this. Data for the study was collected via survey instrument. A total of 203 people in various systems engineering roles in launch vehicle organizations throughout the United States voluntarily participated. Each latent construct of the study was validated using confirmatory factor analysis (CFA). Structural equation modeling (SEM) was used to examine the relationships between the variables of the study. The IBM SPSS Amos 25 software was used to analyze the CFA and SEM.
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Date Issued
2019
Identifier
CFE0007806, ucf:52348
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0007806
INDIA AND CHINA SPACE PROGRAMS: FROM GENESIS OF SPACE TECHNOLOGIES TO MAJOR SPACE PROGRAMS AND WHAT THAT MEANS FOR THE INTERNATIONAL COMMUNITY
Title
INDIA AND CHINA SPACE PROGRAMS: FROM GENESIS OF SPACE TECHNOLOGIES TO MAJOR SPACE PROGRAMS AND WHAT THAT MEANS FOR THE INTERNATIONAL COMMUNITY.
Creator
BHOLA, GAURAV, HANDBERG, ROGER, University of Central Florida
Abstract / Description
The Indian and Chinese space programs have evolved into technologically advanced vehicles of national prestige and international competition for developed nations. The programs continue to evolve with impetus that India and China will have the same space capabilities as the United States with in the coming years. This will present new challenges to the international community in spheres civilian, to space and military applications and their residual benefits.
Date Issued
2009
Identifier
CFE0002745, ucf:48156
Format
Document (PDF)
PURL
http://purl.flvc.org/ucf/fd/CFE0002745