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1
Sellers, Jerry Jon. "Investigation into hybrid rockets and other cost-effective propulsion system options for small satellites." Thesis, University of Surrey, 1986. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.309201.
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Vanherweg, Joseph B. R. "HYBRID ROCKET MOTOR SCALING PROCESS." DigitalCommons@CalPoly, 2015. https://digitalcommons.calpoly.edu/theses/1394.
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Hybrid rocket propulsion technology shows promise for the next generation of sounding rockets and small launch vehicles. This paper seeks to provide details on the process of developing hybrid propulsion systems to the academic and amateur rocket communities to assist in future research and development. Scaling hybrid rocket motors for use in sounding rockets has been a challenge due to the inadequacies in traditional boundary layer analysis. Similarity scaling is an amendment to traditional boundary layer analysis which is helpful in removing some of the past scaling challenges. Maintaining geometric similarity, oxidizer and fuel similarity and mass flow rate to port diameter similarity are the most important scaling parameters. Advances in composite technologies have also increased the performance through weight reduction of sounding rockets through and launch vehicles. Technologies such as Composite Overwrapped Pressure Vessels (COPV) for use as fuel and oxidizer tanks on rockets promise great advantages in flight performance and manufacturing cost. A small scale COPV, carbon fiber ablative nozzle and a N class hybrid rocket motor were developed, manufactured and tested to support the use of these techniques in future sounding rocket development. The COPV exhibited failure within 5% of the predicted pressure and the scale motor testing was useful in identifying a number of improvements needed for future scaling work. The author learned that small scale testing is an essential step in the process of developing hybrid propulsion systems and that ablative nozzle manufacturing techniques are difficult to develop. This project has primarily provided a framework for others to build upon in the quest for a method to easily develop hybrid propulsion systems sounding rockets and launch vehicles.
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Chamberlain, Britany L. "Additively-Manufactured Hybrid Rocket Consumable Structure for CubeSat Propulsion." DigitalCommons@USU, 2018. https://digitalcommons.usu.edu/etd/7285.
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Three-dimensional, additive printing has emerged as an exciting new technology for the design and manufacture of small spacecraft systems. Using 3-D printed thermoplastic materials, hybrid rocket fuel grains can be printed with nearly any cross-sectional shape, and embedded cavities are easily achieved. Applying this technology to print fuel materials directly into a CubeSat frame results in an efficient, cost-effective alternative to existing CubeSat propulsion systems. Different 3-D printed materials and geometries were evaluated for their performance as propellants and as structural elements. Prototype "thrust columns" with embedded fuel ports were printed from a combination of acrylonitrile utadiene styrene (ABS) and VeroClear, a photopolymer substitute for acrylic. Gaseous oxygen was used as the oxidizer for hot-fire testing of prototype thrusters in ambient and vacuum conditions. Hot-fire testing in ambient and vacuum conditions on nine test articles with a combined total of 25 s burn time demonstrated performance repeatability. Vacuum specific impulse was measured at over 167 s and maximum thrust of individual thrust columns at 9.5 N. The expected ΔV to be provided by the four thrust columns of the consumable structure is approximately 37 m/s. With further development and testing, it is expected that the consumable structure has the potential to provide a much-needed propulsive solution within the CubeSat community with further applications for other small satellites.
4
Seubert, Carl Reiner. "Refrigerant-based propulsion system for small spacecraft." Diss., Rolla, Mo. : University of Missouri-Rolla, 2007. http://scholarsmine.umr.edu/thesis/pdf/Carl_Reiner_Seubert_Masters_Thesis_09007dcc8031c34d.pdf.
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Thesis (M.S.)--University of Missouri--Rolla, 2007.Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed May 11, 2007) Includes bibliographical references (p. 115-119).
5
Clough, Joshua. "Integrated propulsion and power modeling for bimodal nuclear thermal rockets." College Park, Md.: University of Maryland, 2007. http://hdl.handle.net/1903/7604.
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Thesis (Ph. D.) -- University of Maryland, College Park, 2007.Thesis research directed by: Dept. of Aerospace Engineering. Title from t.p. of PDF. Includes bibliographical references. Published by UMI Dissertation Services, Ann Arbor, Mich. Also available in paper.
6
Westerlund, Simon. "Design of Ablative Insulator for Solid Rocket Booster." Thesis, KTH, Energiteknik, 2015. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-179031.
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The objective of this master thesis was to investigate an ablative liner for the T-Minus DART booster that will accelerate a dart to Mach 5.2 within five seconds. An oxyacetylene torch test was used to sort out the obviously bad materials. Glass fiber/epoxy, with and without alumina as fire retardant, and carbon fiber/epoxy were selected for further investigation. A sub-scale motor was built to expose the materials for conditions similar to the booster conditions in regard to temperature, chemistry, flow velocity and pressure. The target pressure could not be reached in the sub-scale motor but a polynomial function was fitted to the data in order to extrapolate the data and estimate the ablation rate at 7 MPa. The final design is always based on measurements on full scale motors. This could not be done within this report. Recommendation for future work is to use an insulator of 1.8 mm of carbon fiber/epoxy or 1.3 mm of glass fiber/epoxy/alumina for the sub-scale firings to come.
7
Ziemba, Timothy Martin. "Experimental investigation of the mini-magnetospheric plasma propulsion prototype /." Thesis, Connect to this title online; UW restricted, 2003. http://hdl.handle.net/1773/9962.
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Harper, James M. "Pocket Rocket: A 1U+ Propulsion System Design To Enhance CubeSat Capabilities." DigitalCommons@CalPoly, 2020. https://digitalcommons.calpoly.edu/theses/2218.
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The research presented provides an overview of a 1U+ form factor propulsion system design developed for the Cal Poly CubeSat Laboratory (CPCL). This design utilizes a Radiofrequency Electrothermal Thruster (RFET) called Pocket Rocket that can generate 9.30 m/s of delta-V with argon, and 20.2 ± 3 m/s of delta-V with xenon. Due to the demand for advanced mission capabilities in the CubeSat form factor, a need for micro-propulsion systems that can generate between 1 – 1500 m/s of delta-V are necessary.
By 2019, Pocket Rocket had been developed to a Technology Readiness Level (TRL) of 5 and ground tested in a 1U CubeSat form factor that incorporated propellant storage, pressure regulation, RF power and thruster control, as well as two Pocket Rocket thrusters under vacuum, and showcased a thrust of 2.4 mN at a required 10 Wdc of power with Argon propellant. The design focused on ground testing of the thruster and did not incorporate all necessary components for operation of the thruster. Therefore in 2020, a 1U+ Propulsion Module that incorporates Pocket Rocket, the RF amplification PCB, a propellant tank, propellant regulation and delivery, as well as a DC-RF conversion with a PIB, that are all attached to a 2U customer CubeSat for a 3U+ overall form factor. This design was created to increase the TRL level of Pocket Rocket from 5 to 8 by demonstrating drag compensation in a 400 km orbit with a delta-V of 20 ± 3 m/s in the flight configuration. The 1U+ Propulsion Module design included interface and requirements definition, assembly instructions, Concept of Operations (ConOps), as well as structural and thermal analysis of the system. The 1U+ design enhances the capabilities of Pocket Rocket in a 1U+ form factor propulsion system and increases future mission capabilities as well as propulsion system heritage for the CPCL.
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Lugtu, Spotrizano Descanzo. "Impact of ion propulsion on performance, design, testing and operation of a geosynchronous spacecraft." Thesis, Monterey, California : Naval Postgraduate School, 1990. http://handle.dtic.mil/100.2/ADA237028.
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Thesis (M.S. in Astronautical Engineering)--Naval Postgraduate School, June 1990.Thesis Advisor(s): Agrawal, Brij N. Second Reader: Biblarz, Oscar. "June 2009." Description based on title screen as viewed on 19 October 2009. DTIC Identifier(s): Ion propulsion, synchronous satellites, NSSK (North South Station Keeping). Author(s) subject terms: Ion propulsion, geosynchronous satellite, North-South Station Keeping. Includes bibliographical references (p. 153-156). Also available in print.
10
Vernacchia, Matthew T. "Development, modeling and testing of a slow-burning solid rocket propulsion system." Thesis, Massachusetts Institute of Technology, 2017. http://hdl.handle.net/1721.1/112515.
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Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2017.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 163-168).
Small, unmanned aerial vehicles (UAVs) are expanding the capabilities of aircraft systems. However, a gap exists in the size and capability of aircraft: no aircraft smaller than 10 kilograms are capable of flight faster than 100 meters per second. A small, fast aircraft requires a propulsion system which is both miniature and high-power, requirements which current UAV propulsion technologies do not meet. To meet this need, a slow-burning solid rocket motor has been developed. Such motors require slow-burning solid propellants with tailorable burn rate. This thesis reports experimental results and combustion theory for a slow-burning solid propellant. It also describes a rocket motor designed to use this propellant, and the manufacturing process used to produce it. This propellant burns slowly enough for the low-thrust, long-endurance needs of UAV propulsion. Its burn rate can be predictably tailored by addition of the burn rate suppressant oxamide. Further, this thesis presents a concept for a small, fast aircraft designed around this novel propulsion technology. The motor integrates elegantly into the aircraft's structure, and compact thermal protection system insulates other vehicle systems from the heat of combustion. These results demonstrate the feasibility slow-burning rocket propulsion systems, and their application to small aircraft. It should be possible for small, rocket-propelled UAVs to sustain powered, transonic flight for several minutes. With this technology, kilogram-scale UAVs could be able to quickly deploy over tens of kilometers, and fly joint missions alongside manned fighter jets.
by Matthew T. Vernacchia.
S.M.
11
Vernacchia, Matthew T. "Development of low-thrust solid rocket motors for small, fast aircraft propulsion." Thesis, Massachusetts Institute of Technology, 2020. https://hdl.handle.net/1721.1/127069.
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Thesis: Ph. D., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, May, 2020Cataloged from the official PDF of thesis.
Includes bibliographical references (pages 281-289).
Small, uncrewed aerial vehicles (UAVs) are expanding the capabilities of aircraft systems. However, a gap exists in the size and capability of aircraft: no small aircraft are capable of sustained fast flight. A small, fast aircraft requires a propulsion system which is both miniature and high-power, requirements which current UAV propulsion technologies do not meet. Solid propellant rocket motors could be used, but must be re-engineered to operate at much lower thrust and for much longer burn times than conventional small solid rocket motors. This imposes unique demands on the motor and propellant. This work investigates technological challenges of small, low-thrust solid rocket motors: slow-burn solid propellants, motors which have low thrust relative to their size (and thus have low chamber pressure), thermal protection for the motor case, and small nozzles which can withstand long burn times.
Slow-burn propellants were developed using ammonium perchlorate oxidizer and the burn rate suppressant oxamide. By varying the amount of oxamide (from 0-20%), burn rates from 4mms⁻¹ to 1mms⁻¹ (at 1MPa) were achieved. Using these propellants, a low-thrust motor successfully operated at a (thrust / burn area) ratio 10 times less than that of typical solid rocket motors. This motor can provide 5-10N of thrust for 1-3 minutes. An ablative thermal protection liner was tested in these firings. Despite the long burn time, only a few millimeters of ablative are needed. A new ceramic-insulated nozzle was demonstrated on this motor. The nozzle has a small throat diameter (only a few millimeters) and can operate in thermal steady-state. Models were developed for the propellant burn rate, motor design, heat transfer within the motor and nozzle, and for thermal stresses in the nozzle insulation.
This work shows that small, low-thrust solid motors are feasible, by demonstrating these key technologies in a prototype motor. Further, the experimental results and models will enable engineers to design and predict the performance of solid rocket motors for small, fast aircraft. By providing insight into the physics of these motors, this thesis may help to enable a new option for aircraft propulsion.
by Matthew T. Vernacchia.
Ph. D.
Ph.D. Massachusetts Institute of Technology, Department of Aeronautics and Astronautics
12
Ridard, Mathilde. "NUMERICAL STUDY OF TWO-PHASE FLOWS FOR PERFORMANCES IN SOLID ROCKET PROPULSION." Thesis, KTH, Skolan för industriell teknik och management (ITM), 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-218006.
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Montre, Trevor Allen. "Experimental Investigation of a 2-D Air Augmented Rocket: Effects of Nozzle Lip Thickness on Rocket Mixing and Entrainment." DigitalCommons@CalPoly, 2011. https://digitalcommons.calpoly.edu/theses/662.
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Cold-flow tests were performed using a simulated Air Augmented Rocket (AAR) operating as a mixer-ejector in order to investigate the effects of varied primary nozzle lip thickness on mixing and entrainment. The simulated primary rocket ejector was supplied with nitrogen at a maximum chamber stagnation pressure of 1712 psi, and maximum flow rate of 1.67 lbm/s. Secondary air was entrained from a plenum, producing pressures as low as 6.8 psi and yielding maximum stagnation pressure ratios as high as 160. The primary ejector nozzles each had an area ratio of approximately 20, yielding average primary exit Mach numbers between 4.34 and 4.57. The primary flow was ejected into an 18.75 inch-long mixing duct with a rectangular cross-sectional area of 2.10 in2. The secondary flow was entrained into the mixing duct through a total cross section of 0.94 in2. Two mixing duct configurations were used, one with plexiglass upper and lower surfaces for flow visualization and one with pressure ports along the lower surface for primary plume measurements.
Shadowgraph images were used to characterize the mixing duct flow field, while pressure and temperature instrumentation allowed for calculation of various ejector performance characteristics. Experimentally-calculated performance characteristics were compared to inviscid theoretical predictions. Varying degrees of flow field asymmetry were observed with each nozzle. Test repeatability was found to be excellent for all nozzles. Several distinct phenomena were observed in both the primary plume and secondary streams.
The duration of secondary flow choking was found to be inversely proportional to nozzle lip thickness, due to the primary plume being physically closer to the secondary flow with a thinner nozzle lip. This indicated that the ejector’s ability to choke the secondary flow is primarily an inviscid phenomenon.
Secondary flow blockage was demonstrated in two consecutive tests using the thickest nozzle lip. Only the left secondary duct became blocked in each case. Blockage was only demonstrated in the centerline pressure configuration, so no visual evidence was able to support the blocked flow theory.
At every pressure ratio, entrainment ratio was shown to increase with nozzle lip thickness. The original conical nozzle produced the largest level of entrainment, indicating that the angle of primary flow impingement was the largest contributing factor to secondary entrainment. The increase in efficiency resulting from a bell-mouth nozzle was less than the increase in entrainment efficiency of a conical nozzle, indicating that the conical design was more efficient overall for air augmented rocket applications.
14
Siebert, Joseph R. "Design hazard analysis, and system level testing of a university propulsion system for spacecraft application." Diss., Rolla, Mo. : Missouri University of Science and Technology, 2009. http://scholarsmine.mst.edu/thesis/pdf/Siebert_09007dcc8063c59e.pdf.
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Thesis (M.S.)--Missouri University of Science and Technology, 2009.Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed April 14, 2009) Includes bibliographical references (p. 201-203).
15
Popish, Martin Roy. "Primary and Secondary Flow Interactions in the Mixing Duct of a 2-D Planer Air Augmented Rocket." DigitalCommons@CalPoly, 2012. https://digitalcommons.calpoly.edu/theses/766.
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Experiments were conducted on the Cal Poly air augmented rocket (AAR) in order to characterize two-dimensional flowfield phenomenon occurring in the mixing duct. The testing utilized a direct connect system where high pressure nitrogen is fed into the combustion chamber, to form a primary flow. The high pressure nitrogen is then expanded through a nozzle, with an area ratio of 22 and an exit area of 0.75 in2, up to Mach 4.3. Secondary air is entrained from a plenum chamber which is used to create a lower stagnation pressure for the secondary flow. The two flows mix in a duct that has a cross sectional area of 2.06 in2. The maximum pressure ratio, the ratio of primary to secondary stagnation pressure, achieved during testing was 132. The stagnation pressures of the primary and secondary flows are transient throughout the test. The quasi-steady portion of each run increased with increasing pressure ratio. Pressure and temperature measurements were collected from ten test runs.
Shadowgraph images were taken of the mixing duct during testing in order to image the interactions between the primary and secondary flows. The images show an oblique shock forming in the primary flow. The angle of the shock matches theoretical predictions to within 8.41%. The oblique shock begins at a distance of 1.5 inches downstream of nozzle exit when the AAR is operating in the Fabri choked condition. The images also show the mixing region which forms between the primary and secondary flows. The mixing region represents as much as 25% of the cross-sectional area of the flow field in the mixing duct two inches downstream of the nozzle exit.
An analysis of the secondary Mach number in the mixing duct shows that Fabri choking is occurring during testing. The secondary Mach number decreases as pressure ratio increases, in the Fabri choked condition. The transition to Fabri choking occurs at a pressure ratio of 100, suggesting that this is the pressure ratio of the saturated case.
The shape of the primary plume was compared to results from a 2-D simulation developed to predict the flow field inside the Cal Poly AAR. Although, the simulation is unable to predict the entire flowfield, modifications made it able to predict the velocity of the secondary, entrained, flow within 3.7%. The modified simulation also predicts the that the primary plume will have expanded 98% of its total distance from the centerline of the mixing duct 1.7 inches downstream of the primary nozzle exit.
Pressure data taken along the wall of the mixing duct was used to identify the location of Fabri choking in the mixing duct. Tests showed that Fabri choking is occurring between 1 inch and 2.5 inches downstream of the nozzle exit. The location of Fabri choking moves farther downstream of the nozzle as pressure ratio increases.
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Kokan, Timothy Salim. "Characterizing High-Energy-Density Propellants for Space Propulsion Applications." Diss., Georgia Institute of Technology, 2007. http://hdl.handle.net/1853/14626.
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There exists wide ranging research interest in high-energy-density matter (HEDM) propellants as a potential replacement for existing industry standard fuels for liquid rocket engines. The U.S. Air Force Research Laboratory, the U.S. Army Research Lab, the NASA Marshall Space Flight Center, and the NASA Glenn Research Center each either recently concluded or currently has ongoing programs in the synthesis and development of these potential new propellants.
In order to perform conceptual designs using these new propellants, most conceptual rocket engine powerhead design tools (e.g. NPSS, ROCETS, and REDTOP-2) require several thermophysical properties of a given propellant over a wide range of temperature and pressure. These properties include enthalpy, entropy, density, viscosity, and thermal conductivity. Very little thermophysical property data exists for most of these potential new HEDM propellants. Experimental testing of these properties is both expensive and time consuming and is impractical in a conceptual vehicle design environment.
A new technique for determining these thermophysical properties of potential new rocket engine propellants is presented. The technique uses a combination of three different computational methods to determine these properties. Quantum mechanics and molecular dynamics are used to model new propellants at a molecular level in order to calculate density, enthalpy, and entropy. Additivity methods are used to calculate the kinematic viscosity and thermal conductivity of new propellants.
This new technique is validated via a series of verification experiments of HEDM compounds. Results are provided for two HEDM propellants: quadricyclane and 2-azido-N, N-dimethylethanamine (DMAZ). In each case, the new technique does a better job than the best current computational methods at accurately matching the experimental data of the HEDM compounds of interest.
A case study is provided to help quantify the vehicle level impacts of using HEDM propellants. The case study consists of the National Aeronautics and Space Administrations (NASA) Exploration Systems Architecture Study (ESAS) Lunar Surface Access Module (LSAM). The results of this study show that the use of HEDM propellants instead of hypergolic propellants can lower the gross weight of the LSAM and may be an attractive alternative to the current baseline hypergolic propellant choice.
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Krolak, Matthew Joseph. "Optimization of a magnetoplasmadynamic arc thruster." Link to electronic thesis, 2007. http://www.wpi.edu/Pubs/ETD/Available/etd-042607-155701/.
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Romano, Federico. "Q1D unsteady ballistic model for solid rocket motors performance prediction." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2021.
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The simulation tool ROBOOST, in use at the Alma Propulsion Lab of the University of Bologna – Forlì Campus, exploits a hybrid ballistic model 0D-1D. The need of a complete Q1D model for the entire combustion time, from motor start-up to burn out arised. The present work is devoted to the development and test of a Q1D unsteady ballistic model for solid rocket motors performance prediction. The newly developed code, called SOL1D, is written in Matlab environment and is capable of predicting the time and space evolution of all the main thermodynamic variables during the solid rocket motor combustion process. The model has been tested and validated on a BARIA motor, thus demonstrating its adherence to experimental data. SOL1D paves the way for future works aimed at simulating performances of actual launchers.
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Biddy, Christopher Lorian. "Development of a High Performance Micropropulsion System for CubeSats." DigitalCommons@CalPoly, 2009. https://digitalcommons.calpoly.edu/theses/150.
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Picosatellites are defined as satellites with a mass between 0.1 and 1kg (Miniaturized satellite). Picosatellites are typically designed to work together or function in formations (Miniaturized satellite). A specific type of Picosatellite known as CubeSats were introduced in 1999 and since then have increased in popularity so that there are now over 80 CubeSat programs around the world. CubeSats are defined as cubic units 10cm on each side and no more than 1kg in mass. CubeSats are required to conform to the CubeSat Standard created by California Polytechnic State University and Stanford University and be compatible with Cal Poly’s P-POD deployment system (Toorian, 2005). Some CubeSat uses include earth imaging, communications projects and various scientific experiments. CubeSats currently require attitude control and in the future, may require, maintaining a specific orbit, or changing orbit. With this ability many new activities may be possible for CubeSats. These activities could include rendezvous, vehicle inspection, formation flying and de-orbiting. For these activities to be possible, a high performance propulsion system is required. The goal of this thesis is to design and test an affordable, safe, and effective micro-propulsion system for CubeSats.
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Santos, Genivaldo Pimenta dos. "Experimental evaluation of hybrid propulsion rocket engine operating with paraffin fuel grain and gaseous oxygen." Instituto Tecnológico de Aeronáutica, 2014. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=2967.
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In the last decade the hybrid propulsion has been considering as a viable alternative of chemical energy conversion stored in propellants into kinetic energy. This energy is applied in propulsive systems of manned platforms, maneuvering procedures and even in the repositioning process of micro satellites. It presents attractive features and good balance between performance and environmental impact. Paraffin based grains are the hybrid solid fuels appointed as polymeric fuel substitute. The liquid layer formed on the burning surface ensures high regression rate when driven into the flame front. Paraffin grains allow raw material recovery and reduce the risk of explosion in the presence of erosive burning. The structure of the grain and the control of the liquefying burning surface layer depend on the additives concentration, such as carbon black, which are added to the fuel matrix during the production process. In the solid propellant paraffin based grain a cylindrical center port developed during the centrifugation tends to concentrate carbon black in the outer region of the grain. The influence of carbon black distribution and hardness gradient in paraffin based grain were evaluated in this work. Despite being a well-known material, scarce data on the relation of activation energy (Ea) of paraffin is available. In this work, the kinetic parameters (activation energy and pre-exponential factor) of microcrystalline 140/1450F paraffin have been raised through Thermo Gravimetric Analysis in conjunction with the Arrhenius kinetic mechanism, according to ASTM-E1461. The analysis indicated that the microcrystalline 140/1450F paraffin presents activation energy of 242 kJ.mol-1 and pre-exponential factor from 1.42x1020 min-1 to 2.90x1025 min-1. Ignition was achieved with a 50 W pyrotechnic igniter. Firing tests with 140/1450F paraffin as solid fuel and gaseous oxygen (GOX) as oxidizer were carry on at pressure above 3.8 MPa. The study suggests that multiple thin layers grain may generate burning surfaces with hardness and carbon black concentration almost constant. In this work combustion instability presented by rocket engine was calculated applying the frequency and boundary layer delay time relationship as proposed by Karabeyoglu. The frequency instability up to was evaluated using LabVIEW data acquisition system. In hybrid propulsion the carbon black has been playing a key role in the gas production process on the burning surface, the contribution of carbon black in the combustion instability suppression of hybrid propulsion system with paraffin-based propellant was evaluated. The results confirm the potential use of paraffin base propellant grains loaded with carbon black charge in hybrid propulsion. Paraffin propellants grains doped with carbon black (CB - from 2% to 8%) were burned on the rocket engine test workbench in the Aeronautical Engineering Laboratory / ITA. In the hybrid combustion performance evaluation based on conventional methods approach, it was applied graphical means to visualize the flow path results. Through the flow profile adjustment it was identified a global regression rates values near the practical grain consumption. The model represented graphically the relationships between oxidizer mass flow rate from with injection pressure from and it is recommended as a tool to help the hybrid propulsion performance assessment.
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Schafer, Michael D. "Feasibility of SCRAMJET technology for an intermediate propulsive stage of an expendable launch vehicle." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2002. http://library.nps.navy.mil/uhtbin/hyperion-image/02sep%5FSchafer.pdf.
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Thesis (M.S. in Space Systems Operations)--Naval Postgraduate School, September 2002.Thesis advisor(s): Stephen A. Whitmore, Charles M. Racoosin. Includes bibliographical references (p. 93). Also available online.
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Tucker, J. "A whole life assessment of extruded double base propellants." Thesis, Cranfield University, 2013. http://dspace.lib.cranfield.ac.uk/handle/1826/8032.
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The manufacturing process for solventless extruded double base propellants involves a number of rolling and reworking stages. Throughout these processes a decrease in weight average molecular weight was observed, this was attributed to denitration. Differential scanning calorimetery data indicated that the reworking stages of extruded double base propellant manufacture were crucial to the homogenisation of the propellant mixture. To determine the homogeneity of the final extruded product, a sample was analysed across its diameter. No variations in stabiliser concentration, molecular weight, or Vickers hardness were detected. An accelerated thermal ageing trial simulating up to 8 years of ageing at 25°C was carried out to evaluate the storage characteristics. Reductions in stabiliser concentration, number average molecular weight, weight average molecular weight and polydispersity compared with un-aged samples were observed. The glass transition temperature measured using differential scanning calorimetery decreased by ~3°C. The decrease was attributed to the initial denitration reducing the energy of bond rotation and shortening the polymer chains, both factors reducing the energy required for movement. Modulus values determined from dynamic mechanical analysis temperature scanning experiments, did not detect significant variation between un-aged and aged samples. Though it was considered that variations would be likely if a more extensive ageing program was completed. In order to evaluate propellant behaviour at very high and low frequencies, time temperature superposition (TTS) and creep testing were carried out. The TTS technique superpositioned data well, allowing future investigation of high frequency propellant properties. Creep testing was considered to be an appropriate approach, though the equipment available was not optimised for such testing. This thesis is concerned with understanding how propellants are manufactured from nitrocellulose, nitroglycerine and other constituents. It is also about how the propellants decompose during long periods of time in storage, and how these changes can be measured using thermal and mechanical methods. It is about how the physical, chemical and thermal properties of the propellant composition change throughout the manufacture. This is relevant as it could be used to develop more efficient manufacturing processes, allow operators to adjust processes to tailor product properties or be used to re-design manufacturing to compensate for a different starting material. The thesis also considers how and why the properties of the product change over the course of years of storage. A specific focus on whether changes in mechanical and thermal properties occur, and if so how they can be detected.
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Johnson, Kyle Jacob. "AXISYMMETRIC AIR AUGMENTED METHANOL/GOX ROCKET MIXING DUCT EXPERIMENTAL THRUST STUDY." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/930.
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A hot-flow axisymmetric Air Augmented Rocket (AAR) test apparatus was constructed to test various mixing duct configurations at static conditions. Primary flow for the AAR was provided through a liquid methanol-gaseous oxygen bipropellant rocket. Experimental thrust measurements were recorded and propellant mass flow rates and chamber conditions were calculated using an iterative solver dependant on recorded propellant line stagnation pressures. Primary rocket flow produced thrust ranging from 14 to 17.9lbf. Primary mass flow rate through testing ranged from 0.071 to 0.085lbm/s with calculated chamber pressures between 298-362psia. Calculated primary flow velocity ranged from 6,600ft/s to 8,000ft/s depending on propellant pressure inputs and calculated chamber conditions.
The AAR test apparatus was capable of testing various mixing duct geometries and measuring the axial thrust of the mixing ducts separately from the total thrust of the system. Two mixing duct geometries, a straight wall mixing duct and diverging wall mixing duct, with identical exterior dimensions and inlet geometry were tested for a range of air/fuel mixture ratios from 0.82 to 2.2 spanning the stoichometric mixture ratio of 1.5. Mixing duct thrust did not vary greatly with primary flow characteristics. Straight mixing duct thrust averaged 0.97lbf and diverging mixing duct thrust averaged 0.18lbf. Total system thrust decreased by an average of 0.62lbf with a straight mixing duct and 0.74lbf with a diverging mixing duct. Decreases in total thrust are attributed to low pressure flow interaction between the mixing duct and the primary rocket assembly.
Visual flow comparison between mixing duct configurations and fuel ratio cases were carried out using high definition video recording with a grid reference for comparison. The diverging mixing duct produced the greatest variation in visible flow when compared to a straight mixing duct and no mixing duct configuration. This indicated that the diverging mixing duct had a greater influence on primary and secondary flow field mixing than the straight mixing duct.
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Masquelet, Matthieu Marc. "Large-eddy simulations of high-pressure shear coaxial flows relevant for H2/O2 rocket engines." Diss., Georgia Institute of Technology, 2013. http://hdl.handle.net/1853/47522.
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The understanding and prediction of transient phenomena inside Liquid Rocket Engines
(LREs) have been very difficult because of the many challenges posed by the
conditions inside the combustion chamber. This is especially true for injectors involving
liquid oxygen LOX and gaseous hydrogen GH₂. A wide range of length scales
needs to be captured from high-pressure flame thicknesses of a few microns to the length
of the chamber of the order of a meter. A wide range of time scales needs to be captured,
again from the very small timescales involved in hydrogen chemistry to low-frequency
longitudinal acoustics in the chamber. A wide range of densities needs to be captured,
from the cryogenic liquid oxygen to the very hot and light combustion products. A wide
range of flow speeds needs to be captured, from the incompressible liquid oxygen jet to
the supersonic nozzle. Whether one desires to study these issues numerically or
experimentally, they combine to make simulations and measurements very difficult whereas
reliable and accurate data are required to understand the complex physics at stake. This
thesis focuses on the numerical simulations of flows relevant to LRE applications
using Large Eddy Simulations (LES). It identifies the required features to tackle
such complex flows, implements and develops state-of-the-art solutions
and apply them to a variety of increasingly difficult problems.
More precisely, a multi-species real gas framework is developed inside a conservative,
compressible solver that uses a state-of-the-art hybrid scheme to capture at the same time
the large density gradients and the turbulent structures that can be found in a
high-pressure liquid rocket engine.
Particular care is applied to the
implementation of the real gas framework with detailed derivations of thermodynamic
properties, a modular implementation of select equations of state in the solver.
and a new efficient iterative method.
Several verification cases are performed to evaluate this implementation and the
conservative properties of the solver. It is then validated against laboratory-scaled
flows relevant to rocket engines, from a gas-gas reacting injector to a liquid-gas
injector under non-reacting and reacting conditions. All the injectors considered contain
a single shear coaxial element and the reacting cases only deal with H₂-O₂ systems.
A gaseous oyxgen-gaseous hydrogen (GOX-GH₂) shear coaxial injector, typical
of a staged combustion engine, is first investigated. Available experimental data is
limited to the wall heat flux but extensive comparisons are conducted between
three-dimensional and axisymmetric solutions generated by this solver as well as by other
state-of-the-art solvers through a NASA validation campaign. It is found that the unsteady
and three-dimensional character of LES is critical in capturing physical flow features,
even on a relatively coarse grid and using a 7-step mechanism instead of a 21-step
mechanism. The predictions of the wall heat flux, the only available data, are not very good and
highlight the importance of grid resolution and near-wall models for LES.
To perform more quantitative comparisons, a new experimental setup is investigated under
both non-reacting and reacting conditions. The main difference with the previous setup,
and in fact with most of the other laboratory rigs from the literature, is the presence of
a strong co-flow to mimic the surrounding flow of other injecting elements. For the
non-reacting case, agreement with the experimental high-speed visualization is very good,
both qualitatively and quantitatively but for the reacting case, only poor agreement is
obtained, with the numerical flame significantly shorter than the observed one. In both
cases, the role of the co-flow and inlet conditions are investigated and highlighted.
A validated LES solver should be able to go beyond some experimental
constraints and help define the
next direction of investigation. For the non-reacting case, a new scaling law is suggested after a
review of the existing literature and a new numerical experiment agrees with the
prediction of this scaling law.
A slightly modified version of this non-reacting setup is
also used to investigate and validate the Linear-Eddy Model (LEM), an advanced sub-grid closure
model, in real gas flows for the first time.
Finally, the structure of the trans-critical
flame observed in the reacting case hints at the need for such more advanced
turbulent combustion model for this class of flow.
25
Massman, Jeffrey. "NUMERICAL FLOW FIELD ANALYSIS OF AN AIR AUGMENTED ROCKET USING THE AXISYMMETRIC METHOD OF CHARACTERISTICS." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/1141.
Full textAbstract:
An Axisymmetric Rocket Ejector Simulation (ARES) was developed to numerically analyze various configurations of an air augmented rocket. Primary and secondary flow field visualizations are presented and performance predictions are tabulated. A parametric study on ejector geometry is obtained following a validation of the flow fields and performance values.
The primary flow is calculated using a quasi-2D, irrotational Method of Characteristics and the secondary flow is found using isentropic relations. Primary calculations begin at the throat and extend through the nozzle to the location of the first Mach Disk. Combustion properties are tabulated before analysis to allow for propellant property selection. Secondary flow calculations employ the previously calculated plume boundary and ejector geometry to form an isentropic solution. Primary and secondary flow computations are iterated along the new pressure distributions established by the 1D analysis until a convergence tolerance is met. Thrust augmentation and Specific Impulse values are predicted using a control volume approach.
For the validation test cases, the nozzle characteristic net is very similar to that of previous research. Plume characteristics are in good agreement but fluctuate in accuracy due to flow structure formulation. The individual unit processes utilized by the Method of Characteristics are found to vary their outputs by up to 0.025% when compared to existing sources. Rocket thrust and specific impulse are increased by up to 22% for a static system and 15% for an ejector flow at Mach 0.5. Evidence of Fabri conditions were observed in the flow visualization and graphically through the performance predictions. It was determined that the optimum ejector divergence angle for an air augmented rocket greatly depends on the stagnation pressure ratio between the primary and secondary flows.
26
Avenall, Ryan Jeffrey. "Use of metallic foams for heat-transfer enhancement in the cooling jacket of a rocket propulsion element." [Gainesville, Fla.] : University of Florida, 2004. http://purl.fcla.edu/fcla/etd/UFE0008720.
Full text
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Capatina, Allen A. C. "AXISYMMETRIC BI-PROPELLANT AIR AUGMENTED ROCKET TESTING WITH ANNULAR CAVITY MIXING ENHANCEMENT." DigitalCommons@CalPoly, 2015. https://digitalcommons.calpoly.edu/theses/1493.
Full textAbstract:
Performance characterization was undertaken for an air augmented rocket mixing duct with annular cavity configurations intended to produce thrust augmentation. Three mixing duct geometries and a fully annular cavity at the exit of the nozzle were tested to enable thrust comparisons. The rocket engine used liquid ethanol and gaseous oxygen, and was instrumented with sensors to output total thrust, mixing duct thrust, combustion chamber pressure, and propellant differential pressures across Venturi flow measurement tubes.
The rocket engine was tested to thrust maximum, with three different mixing ducts, three major combustion pressure sets, and a nozzle exit plane annular cavity (a grooved ring). The combustion pressures tested were , , and allowing for a nozzle pressure ratio range of relative to ambient pressure. The mixture ratio was fuel rich throughout all tests. The engine operated very consistently throughout all the tests performed; however, pressure losses in the feed system prevented higher combustion pressures from being tested.
Three mixing ducts of the same outer diameter were tested. The short and diverging ducts were the same length and the long duct was long. The short and long ducts created positive mixing duct thrust and the diverging duct created negative mixing duct thrust. The long duct case did show better performance than the no duct case when the total thrust was divided by combustion pressure and nozzle throat area. The long duct always created several times more mixing duct thrust than either the short or diverging ducts, but none of the mixing ducts created positive overall thrust augmentation in the over expanded cases tested. The mixing duct thrusts ranged between and . As the combustion pressures were increased, getting closer the nozzle’s optimal expansion, the mixing duct thrusts started converging indicating a difference between nozzle operation at over expanded and under expanded.
The annular cavity had a noticeable effect on the thrust of the engine and the appearance of the plume. The total thrust of the system was decreased by a maximum of and the plume was more sharply defined when the annular cavity was attached. Better mixing between the primary (engine exhaust) flow and the secondary (ambient air) flow was promoted by the annular cavity because it increased the shear layer’s turbulence and the increased turbulence reduced thrust. The greater mixing also allowed for secondary combustion which made the plumes more sharply defined. The annular cavity was also seen to enhance the mixing duct thrusts for all three mixing ducts.
28
Eilers, Shannon Dean. "Development of the Multiple Use Plug Hybrid for Nanosats (Muphyn) Miniature Thruster." DigitalCommons@USU, 2013. https://digitalcommons.usu.edu/etd/1726.
Full textAbstract:
The Multiple Use Plug Hybrid for Nanosats (MUPHyN) prototype thruster incorporates solutions to several major challenges that have traditionally limited the deployment of chemical propulsion systems on small spacecraft. The MUPHyN thruster offers several features that are uniquely suited for small satellite applications. These features include 1) a non-explosive ignition system, 2) non-mechanical thrust vectoring using secondary fluid injection on an aerospike nozzle cooled with the oxidizer flow, 3) a non-toxic, chemically-stable combination of liquid and inert solid propellants, 4) a compact form factor enabled by the direct digital manufacture of the inert solid fuel grain. Hybrid rocket motors provide significant safety and reliability advantages over both solid composite and liquid propulsion systems; however, hybrid motors have found only limited use on operational vehicles due to 1) difficulty in modeling the fuel flow rate 2) poor volumetric efficiency and/or form factor 3) significantly lower fuel flow rates than solid rocket motors 4) difficulty in obtaining high combustion efficiencies. The features of the MUPHyN thruster are designed to offset and/or overcome these shortcomings. The MUPHyN motor design represents a convergence of technologies, including hybrid rocket regression rate modeling, aerospike secondary injection thrust vectoring, multiphase injector modeling, non-pyrotechnic ignition, and nitrous oxide regenerative cooling that address the traditional challenges that limit the use of hybrid rocket motors and aerospike nozzles. This synthesis of technologies is unique to the MUPHyN thruster design and no comparable work has been published in the open literature.
29
Nelson, Lauren May. "Rayleigh Flow of Two-Phase Nitrous Oxide as a Hybrid Rocket Nozzle Coolant." DigitalCommons@CalPoly, 2009. https://digitalcommons.calpoly.edu/theses/284.
Full textAbstract:
The Mechanical Engineering Department at California Polytechnic State University in San Luis Obispo currently maintains a lab-scale hybrid rocket motor for which nitrous oxide is utilized as the oxidizer in the combustion system. Because of its availability, the same two-phase (gas and liquid) nitrous oxide that is used in the combustion system is also routed around the throat of the hybrid rocket’s converging-diverging nozzle as a coolant. While this coolant system has proven effective empirically in previous tests, the physics behind the flow of the two-phase mixture is largely unexplained. This thesis provides a method for predicting some of its behavior by modeling it using the classic gas dynamics scenarios of Rayleigh and Fanno flows which refer to one-dimensional, compressible, inviscid flow in a constant area duct with heat addition and friction. The two-phase model produced utilizes a separated phase with interface exchange model for predicting whether or not dryout occurs. The Shah correlation is used to predict heat transfer coefficients in the nucleate boiling regime. The homogeneous flow model is utilized to predict pressure drop. It is proposed that a Dittus-Boelter based correlation much like that of Groeneveld be developed for modeling heat transfer coefficients upon the collection of sufficient data.
Data was collected from a series of tests on the hybrid rocket nozzle to validate this model. The tests were first run for the simplified case of an ideal gas (helium) coolant to verify the experimental setup and promote confidence in subsequent two-phase experimental results. The results of these tests showed good agreement with a combined Rayleigh-Fanno model with a few exceptions including: (1) reduced experimental gas pressure and temperature in the annulus entrance and exit regions compared to the model and (2) reduced experimentally measured copper temperatures uniformly through the annulus. These discrepancies are likely explained by the geometry of the flowpath and location of the copper thermocouples respectively. Next, a series of two-phase cooled experiments were run. Similar trends were seen to the helium experiment with regards to entrance and exit regions. The two-phase Rayleigh homogeneous flow model underpredicted pressure drop presumably due to the inviscid assumption. Ambiguity was observed in the fluid temperature measurements but the trend seemed to suggest that mild thermal non-equilibrium existed. In both cases, the dryout model predicted that mist flow (a post-CHF regime) occurred over most of the annulus.
Several modifications should be implemented in future endeavors. These include: (1) collecting more data to produce a heat transfer coefficient correlation specific to the nitrous oxide system of interest, (2) accounting for thermal non-equilibrium, (3) accounting for entrance and exit effects, and (4) developing a two-phase Fanno model.
30
Krise, Jeffrey Raymond. "Characterization of High Inlet Diffusion Low Flow Coefficient Inducer Pumps for Space Propulsion in the Presence of a Cavitation Control Device." BYU ScholarsArchive, 2011. https://scholarsarchive.byu.edu/etd/6212.
Full textAbstract:
Historically inducer pumps have been designed with low inlet diffusion that allows for a gradual pressure rise through the machine that has the ability to slowly collapse any cavitation bubbles that may be present. A novel cavitation control device has been developed by researchers at ConceptsNREC that has been shown in previous experimental work to greatly improve the suction performance of a traditionally designed machine. Computational fluid dynamics (CFD) has been employed to understand the effectiveness of the cavitation control device (CCD) at controlling the conditions that lead to cavitation inception and to determine the impact that the CCD has on the flow. Also the upper limit of design incidence ratio where the CCD is no longer able to control the factors that lead to cavitation inception was to be determined through the CFD approach. All machine geometries and test data were provided by researchers at ConceptsNREC. Two cases were selected for validation work and 32 additional designs were employed in a parametric study where the flow coefficient and design incidence ratio were varied over a typical range of interest for a turbopump application. The results of this computational work show that the CCD is able to control the factors that lead to early cavitation inception. The research shows that the addition of the CCD has an overall stabilizing affect on the flow by significantly decreasing the incidence at the leading edge of the blade. It has been determined that the maximum design incidence ratio where the CCD is able to effectively control the factors that lead to cavitation inception is dependent on the flow coefficient and in general the maximum design incidence ratio decreases as the flow coefficient is increased.
31
Armstrong, Isaac W. "Development and Testing of Additively Manufactured Aerospike Nozzles for Small Satellite Propulsion." DigitalCommons@USU, 2019. https://digitalcommons.usu.edu/etd/7428.
Full textAbstract:
Automatic altitude compensation has been a holy grail of rocket propulsion for decades. Current state-of-the-art bell nozzles see large performance decreases at low altitudes, limiting rocket designs, shrinking payloads, and overall increasing costs. Aerospike nozzles are an old idea from the 1960’s that provide superior altitude-compensating performance and enhanced performance in vacuum, but have survivability issues that have stopped their application in satellite propulsion systems. A growing need for CubeSat propulsion systems provides the impetus to study aerospike nozzles in this application. This study built two aerospike nozzles using modern 3D metal printing techniques to test aerospikes at a size small enough to be potentially used on a CubeSat. Results indicated promising in-space performance, but further testing to determine thermal limits is deemed necessary.
32
Girardello, Carlo. "Optical Analysis of Plasma : Flame Emission in Cryogenic Rocket Engines." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-76097.
Full textAbstract:
This thesis contains the results of optical flame emission measurements of the Vulcain 2.1engine and the plasma emission spectroscopy of the Lumen Project engine. The plume spectroscopyis analyzed, ordered and studied in detail to offer the best possible molecular composition.The main focus relied on the hydroxide radical, blue radiation and other moleculesanalysis of the intensities encountered during the tests. The plasma emission spectroscopy isfocused on the determination of the plasma temperature value in LIBS measurements. Thehydrogen plasma temperature determination of the local thermodynamic equilibrium, followedby the carbon and sequentially oxygen plasma is obtained. The quality of the LTE isto be determined to judge the truthworthness of the determined temperatures. Both the testsare analyzed thanks to the use of spectrographs, cameras and dedicated software for opticalapplications. The results related to the Vulcain 2.1 LOX/LH2 engine showed the evolutionof the plume in different ROF or pressure variations. Furthermore, the results of the LumenProject LOX/methane engine led to the determination of the plasma temperatures and a firstestimation of the LTE quality.Die vorliegende Arbeit präsentiert die Ergebnisse der Abgasstrahlspektroskopie des H2/LOXVulcain 2.1 Triebwerks und der Zündplasma Spektroskopie des CH4/LOX Triebwerks desLUMEN Projektes. Die Abgasstrahlspektroskopie wurde analysiert und im Detail untersuchtum die am besten passende molekulare Zusammensetzung herauszuarbeiten. DasHauptaugenmerk liegt dabei auf dem Hydroxyl- Radikal, der Blauen Strahlung und molekularerIntensitätsanalyse. Bei der Zündplasmaanalyse liegt der Fokus auf der Bestimmungdes LTE Zustands (Lokales thermodynamisches Gleichgewicht) in LIBS. Die Temperaturdes Wasserstoff-, Kohlenstoff und Sauerstoffplasmas wird herangezogen, um die Qualitätdes LTE Zustands zu beurteilen. Für die Testdurchführung wurden Spektrographen, Kamerasund bestimmte Auswertungstools für optische Anwendungen benutzt. Das Verhaltendes Vulcain 2.1 Abgasstrahls abhängig von verschiedenen ROF und Druckstufen ist in denErgebnissen beschrieben. Für das LUMEN Triebwerk konnten erste Zündplasmatemperaturenbestimmt werden und geben einen Rückschluss auf die Qualität des LTE.
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Grieb, Daniel Joseph. "Design and Analysis of a Reusable N2O-Cooled Aerospike Nozzle for Labscale Hybrid Rocket Motor Testing." DigitalCommons@CalPoly, 2012. https://digitalcommons.calpoly.edu/theses/692.
Full textAbstract:
A reusable oxidizer-cooled annular aerospike nozzle was designed for testing on a labscale PMMA-N20[1] hybrid rocket motor at Cal Poly-SLO.[2] The detailed design was based on the results of previous research involving cold-flow testing of annular aerospike nozzles and hot-flow testing of oxidizer-cooled converging-diverging nozzles. In the design, nitrous oxide is routed to the aerospike through a tube that runs up the middle of the combustion chamber. The solid fuel is arranged in an annular configuration, with a solid cylinder of fuel in the center of the combustion chamber and a hollow cylinder of fuel lining the circumference of the combustion chamber. The center fuel grain insulates the coolant from the heat of the combustion chamber. The two-phase mixture of nitrous oxide then is routed through channels that cool the copper surface of the aerospike. The outer copper shell is brazed to a stainless steel core that provides structural rigidity. The gaseous N2O flows from the end of aerospike to provide base bleed, compensating for the necessary truncation of the spike. Sequential and fully-coupled thermal-mechanical finite element models developed in Abaqus CAE were used to analyze the design of the cooled aerospike. The stress and temperature distributions in the aerospike were predicted for a 10-sec burn time of the hybrid rocket motor.
[1] PMMA stands for polymethyl methacrylate, a thermoplastic commonly known by the brand name Plexiglas®. N2O is the molecular formula for nitrous oxide.
[2] California Polytechnic State University, San Luis Obispo
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Sanchez, Josef S. "EXPERIMENTAL INVESTIGATION OF A 2-D AIR AUGMENTED ROCKET: HIGH PRESSURE RATIO AND TRANSIENT FLOW-FIELDS." DigitalCommons@CalPoly, 2012. https://digitalcommons.calpoly.edu/theses/690.
Full textAbstract:
A 2-D Air Augmented Rocket, the Cal Poly Air Augmented Rocket (CPAAR) Test Apparatus operating as a mixer-ejector was tested to investigate high stagnation pressure ratio and transient flow fields of an ejector. The primary rocket ejector was supplied with high pressure nitrogen at a maximum chamber pressure of 1758 psia and a maximum mass flow rate of 1.4 lb/s. The secondary flow air was entrained from a fixed volume plenum chamber producing pressures as low as 3.3 psia. The maximum total pressure ratio achieved was 221. The original CPAAR apparatus was rebuilt re-instrumented and capability expanded. A fixed volume plenum was attached to the secondary ducts through a constant area square section to mimic the cross section of the secondary ducts with a bell mouth inlet. The mixing duct length was increased from 8 in. to 18 in.
An investigation of the mixing duct flow-field was done with data from pressure and temperature instrumentation. A study of the transient operation of the rocket was compared with results from former research to qualify the quasi-steady assumption of the flow-field. The CPAAR produced Fabri-choked operation, the startup transient observed caused the secondary flow to become established during Fabri-choke mode operation. The supersonic saturated mode was not observed during quasi-steady operation. The quasi-steady operation was defined based on characteristics from previous quasi-steady models of transient operation of supersonic ejectors.
The measurement of the data during testing resulted in a 2.96% experimental uncertainty in the entrainment ratio calculation. The smallest entrainment ratio observed was 0.05 at a total pressure ratio of 220. The location of the Fabri-choke point was shown through the interpretation of the primary and secondary flow as a result of the pressure and temperature measurements. The experimental evidence showed the location of the secondary choke point has a logarithmic relationship with the total pressure ratio. At a total pressure ratio of 220, the area of the aerodynamic throat of the secondary flow is 0.26 in2 and the location occurs 6 inches downstream from the nozzle exit. The secondary flow un-choke is related to the breakdown of the shock structure of the primary flow and produces a flow-field asymmetry which blocks the right duct flow.
The CPSE simulation was unable to accurately predict AAR performance when the inputs are changed from the original CPAAR configuration. At high pressure ratios (PR=220), the error in the prediction is 90%.
35
Weyer, Robert Bernhard. "Investigation of the functioning of a liquefied-gas micro-satellite propulsion system." Thesis, Stellenbosch : Stellenbosch University, 2003. http://hdl.handle.net/10019.1/49765.
Full textAbstract:
Thesis (MScEng)--University of Stellenbosch, 2003.ENGLISH ABSTRACT: The focus of this thesis is on the investigation of the functioning of a liquefied-gas thruster. Such a thruster could be used to provide secondary propulsion to a microsatellite in orbit. A general overview of the need for thrusters in micro-satellites is put forward in the introduction. Motivation for deciding to investigate a liquefied-gas system is presented. Recent developments in the field of micro-satellites are discussed as well as their relevance to the project undertaken. Fundamental background theory relevant to the engineering problems associated with the development and analysis of such a system is also presented. Computer programs were written to simulate such a liquefied-gas thruster system. The experimental work carried out to analyse the system from a practical view-point is documented. Attention is also given to the measurement and calibration techniques used to obtain experimental data. One-dimensional fully explicit transient mathematical models of the thruster system were developed to model the system using both compressed air and butane as propellants. These models were incorporated into computer programs used to simulate the transient behaviour of the system. Although it is intended to use butane as the propellant onboard a satellite, the reason for modelling and simulating a system using compressed air is because air is a convenient fluid to work with from both a theoretical and practical point of view. An experimental model of a thruster system was designed, built and tested using air and butane as propellants. Most of the model was built using perspex to allow for the observation of the two-phase behaviour of the propellant inside the system. Locally purchased components were used for the solenoid and fill valves. Readily available butane lighter fluid was used for butane testing. Self-made heating elements were used to provide heat input to the propellant. Testing was done at different back pressures ranging from 100 kPa down to 20 kPa in a vacuum chamber. Good comparison between theoretical and experimental results was obtained for air. Theoretical results for peak thrusts tended to over predict experimental results by approximately 15 % for a system exhausting to a pressure of 100 kPa. Peak thrusts as high as 0.2 N were obtained for vacuum tests conducted at an absolute pressure of 20 kPa. Peak thrusts of approximately 50 mN were achieved for experimental testing III atmospheric conditions using butane with a starting pressure of between 270 and 290 kPa. Typical average thrusts of between 20 mN and 30 mN were noted for butane testing with initial pressure of between 200 to 300 kPa. Peak thrusts of over 0.1 N were observed for vacuum testing at an absolute pressure of 20 kPa. An equation to correlate the experimentally determined average thrust as a function of pulse duration and starting pressure was developed. This correlated most of the experimental data to within ±25 %. Theoretical results for butane testing are able to predict peak thrusts within approximately 20 % for starting pressures in the range of 200 to 300 kPa. Since the project was an exploratory investigation into a liquefied-gas thruster, some additional aspects relating to such systems were also given attention. The effect of liquid propellant motion or sloshing was considered and recommendations regarding the design and placement of the propellant tanks were made. The use of heat pipes as an alternative to electrical heating elements was investigated and some elementary design aspects are presented graphically. The management of the liquid propellant using surface tension devices was examined qualitatively. Recommendations relating to future projects in the field of simple, low-cost propulsion systems for micro-satellites are put forward. More specifically these recommendations are with regard to: thermo-fluid modelling of the propellant, future experimental work to be done, techniques to measure small thrusts and vacuum chamber testing.
AFRIKAANSE OPSOMMING: Die tesis ondersoek die funksionering van 'n vervloeidegas stuwer. So 'n stuwer kan gebruik word om sekondêre aandrywing aan 'n mikro-satelliet in 'n wentelbaan te verskaf. 'n Algemene oorsig oor die behoeftes van stuwers vir mikro-satelliete word voortgesit in die inleiding. Redes vir die gebruik van 'n vervloeidegas stuwer word bespreek. Onlangse ontwikkelinge in die veld van mikro-satelliet aandrywing word bespreek asook die toepaslikheid daarvan. Fundamentele teoretiese agtergrond verbonde aan die ontwikkeling en analise van so 'n stuwer stelsel word ook gegee. Rekenaarprogramme is geskryf om die gedrag van so 'n stuwer stelsel te simuleer. Eksperimentele werk is gedoen om die stelsel vanuit 'n praktiese oogpunt te analiseer. Aandag word ook gegee aan die metings- en kalibrasietegnieke soos toegepas vir die eksperimentele werk. Eendimensionele volle eksplisiete wiskundige modelle is ontwikkelom die oorgangsgedrag van die stuwer-stelsel te simuleer met beide lug en butaan as dryfmiddel. Hierdie modelle is geïnkorporeer in die rekenaar programme om die stuwer stelsel te simuleer. Alhoewel dit beoog word om butaan as die dryfmiddel aan boord die satelliet te gebruik, is lug ook gebruik vir simulasie weens sy gerieflikheid as 'n vloeier uit beide 'n teoretiese en 'n praktiese oogpunt. 'n Eksperimentele model van die stuwer stelsel is ontwerp, gebou en getoets met beide lug en butaan as dryfmiddels. Die model is hoofsaaklik uit perspex gebou sodat die twee-fase gedrag van die butaan uitgebeeld kon word. Vrylik beskikbare butaan aansteker vloeistof IS gebruik VIr butaan toetsing. Selfvervaardigde verhittingselemente is gebruik om hitte aan die dryfmiddel te verskaf. Toetse is gedoen deur verskeie omgewingsdrukke varieërend van 100 kPa af tot 20 kPa in 'n vakuumtenk te gebruik. Goeie ooreenstemming tussen die teoretiese en eksperimentele resultate vir die toetsing van lug is verkry. Die teoretiese resultate neig om die piek stukrag 15 % hoër te voorspel as die eksperimentele resultate vir 'n stelsel wat tot 'n omgewingsdruk van 100 kPa by die uitlaat. Piek stukragte van meer as 0.2 N is gekry vir vakuum toetse wat gedoen is by 'n omgewingsdruk van 20 kPa. Tydens eksperimentele toetsing met butaan teen 'n aanvanklike druk tussen 270 en 290 kPa, in atmosferiese toestande, is piek stukragte van ongeveer 50 mN behaal. Tipiese gemiddelde stukragte van tussen 20 en 30 mN is waargeneem vir butaan toetsing teen 'n aanvanklike druk tussen 200 en 300 kPa. Piek stukragte van meer as 0.1 N is behaal vir vakuum toetse met 'n absolute druk van 20 kPa. 'n Vergelyking om die gemiddelde stukrag, wat eksperimenteel bepaal is, as 'n funksie van puls tydsduur en aanvanklike druk te korreleer, is ontwikkel. Die meeste eksperimentele data se afwyking van die korrelasie-vergelyking was minder as 25 %. Teoretiese resultate vir butaantoetse het piek stukragte binne 20 % van die eksperimenteel metings korrek voorspel vir aanvanklike drukke tussen 200 tot 300 kPa. Weens die feit dat die projek 'n oorhoofse ondersoek in In vervloeidegas stuwer behels, is aandag ook gegee aan addisionele aspekte wat verband hou met sulke stelsels. Die effek van die vloeistof-dryfmiddel se onstabiele beweging in sy tenke is in ag geneem en voorstelle vir die ontwerp en plasing van die dryfmiddel tenke is gemaak. Die gebruik van hitte pype as 'n alternatief vir elektriese verhittingselemente is ondersoek. Verskeie ontwerp aspekte word grafies voorgestel. Die bestuur van die vloeistof-dryfmiddel deur van oppervlak spannings apparaat gebruik te maak, is kwalitatief ondersoek. Voorstelle vir verdere navorsing in die veld van eenvoudige, lae-koste stuwer stelsels vir mikro-satelliete is gemaak. Meer spesifiek is hierdie voorstelle gerig op die termo-vloeidinamiese modellering van die dryfmiddel, verdere eksperimentele navorsing, tegnieke om klein stukragte te meet en vakuumtenk toetse.
36
King, Harrison Raymond. "Electrode Geometry Effects in an Electrothermal Plasma Microthruster." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1899.
Full textAbstract:
Nanosatellites, such as Cubesats, are a rapidly growing sector of the space industry. Their popularity stems from their low development cost, short development cycle, and the widespread availability of COTS subsystems. Budget-conscious spacecraft designers are working to expand the range of missions that can be accomplished with nanosatellites, and a key area of development fueling this expansion is the creation of micropropulsion systems. One such system, originally developed at the Australian National University (ANU), is an electrothermal plasma thruster known as Pocket Rocket (PR). This device heats neutral propellant gas by exposing it to a Capacitively Coupled Plasma (CCP), then expels the heated gas to produce thrust. Significant work has gone towards understanding how PR creates and sustains a plasma and how this plasma heats the neutral gas. However, no research has been published on varying in the device's geometry. This thesis aims to observe how the size of the RF electrode affects PR operation, and to determine if it can be adjusted to improve performance. To this end, a thruster has been built which allows the geometry of the RF electrode to be easily varied. Measurements of the plasma density at the exit of this thruster with different sizes of electrode were then used to validate a Computational Fluid Dynamics (CFD) model capable of approximately reproducing experimental measurements from both this study and from the ANU team. From this CFD, the number of argon ions in the thruster was found for each geometry, since collisions between argon ions and neutrals are primarily responsible for the heating observed in the thruster. A geometry using a 10.5 mm electrode was observed to produce a 23% increase in the quantity of ions produced compared to the baseline 5 mm electrode size, and a 3.5 mm electrode appears to produce 88% more ions.
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Sporschill, Gustave. "Numerical approach of a hybrid rocket engine behaviour : Modelling the liquid oxidizer injection using a Lagrangian solver." Thesis, KTH, Mekanik, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-217231.
Full textAbstract:
To access and operate in space, a wide range of propulsion systems has been developed, from high-thrust chemical propulsion to low-thrust electrical propulsion, and new kind of systems are considered, such as solar sails and nuclear propulsion. Recently, interest in hybrid rocket engines has been renewed due to their attractive features (safe, cheap, flexible) and they are now investigated and developed by research laboratories such as ONERA.This master’s thesis work is in line with their development at ONERA and aims at finding a methodology to study numerically the liquid oxidizer injection using a Lagrangian solver for the liquid phase. For this reason, it first introduces a model for liquid atomiser developed for aeronautical applications, the FIMUR model, and then focuses on its application to a hybrid rocket engine configuration.The FIMUR model and the Sparte solver have proven to work fine with high mass flow rates on coarse grids. The rocket engine simulations have pointed out the need of an initialisation of the flow field. The methodology study has proven that starting with a reduced liquid mass flow rate is preferable to a simulation with a reduced relaxation between the coupled solvers. The former could not be brought to conclusion due to lack of time but gives an encouraging path to further investigate.
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Pérez, Roca Sergio. "Model-based robust transient control of reusable liquid-propellant rocket engines." Thesis, université Paris-Saclay, 2020. http://www.theses.fr/2020UPASS017.
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La tendance actuelle vers un accès plus abordable à l'espace se traduit par des lanceurs et moteurs réutilisables. Du point de vue du contrôle, ces moteurs fusée à propergol liquide (MFPL) réutilisables impliquent des spécifications de robustesse plus exigeantes que ceux à usage unique, principalement en raison de leurs capacités de redémarrage multiple et de modulation de poussée. Classiquement, le système de contrôle gère les opérations des MFPL autour d'un ensemble fini de points prédéfinis. Cette approche réduit leur domaine de modulation à un intervalle restreint dans lequel ils sont conçus pour être sûrs. De plus, les phases transitoires, qui ont un impact important sur la vie du moteur, ne sont pas exécutées de manière robuste. L’objectif de ce travail est donc de développer une boucle de régulation adaptée à l’ensemble des phases d'opération (transitoire et régime permanent) et robuste aux variations paramétriques internes. Plusieurs blocs ont été développés pour constituer la boucle de régulation : simulation de moteur, génération de référence et contrôleurs. Des simulateurs représentatifs des moteurs à cycle générateur de gaz ont tout d'abord été construits. La modélisation purement thermodynamique du cycle a ensuite été adaptée au contrôle, afin d'obtenir des modèles non-linéaires sous forme d'état. Dans ces modèles, l'influence des entrées de commande continues (ouvertures des vannes) et des entrées discrètes (activation des allumeurs et démarreur) est considérée dans un cadre hybride simplifié. La sous-phase continue du transitoire de démarrage est contrôlée en boucle fermée pour suivre des trajectoires de référence pré-calculées. Outre le démarrage, les scénarios de modulation présentent également un algorithme pour le suivi des états finaux. Une méthode de contrôle à base de modèles, la commande prédictive, a été appliquée de manière linéarisée avec des considérations de robustesse à tous ces scénarios, dans lesquels des contraintes dures doivent être respectées. Le suivi des points de fonctionnement en pression (poussée) et du rapport de mélange dans l'enveloppe de conception est atteint en simulation tout en respectant les contraintes. La robustesse aux variations des paramètres, qui sont identifiés comme prédominants par des analyses, est également démontrée. Ce travail ouvre la voie à la validation expérimentale par des simulations hardware-in-the-loop ou des tests sur banc d'essaiThe current trend towards a more affordable access to space is materialising in reusable launchers and engines. From the control perspective, these reusable liquid-propellant rocket engines (LPRE) imply more demanding robustness requirements than expendable ones, mainly due to their multi-restart and thrust-modulation capabilities. Classically, the control system handles LPRE operation at a finite set of predefined points. That approach reduces their throttability domain to a narrow interval in which they are designed to be safe. Moreover, transient phases, which have a great impact on engine life, are not robustly operated. Hence, the goal of this work is to develop a control loop which is adapted to the whole set of operating phases, transient and steady-state, and which is robust to internal parametric variations. Several blocks have been developed to constitute the control loop: engine simulation, reference generation and controllers. First, simulators representative of the gas-generator-cycle engines were built. The purely thermo-fluid-dynamic modelling of the cycle was subsequently adapted to control, obtaining nonlinear state-space models. In these models, the influence of continuous control inputs (valve openings) and of discrete ones (igniters and starter activations) is considered within a simplified hybrid approach. The continuous sub-phase of the start-up transient is feedback controlled to track pre-computed reference trajectories. Beyond the start-up, throttling scenarios also present an end-state-tracking algorithm. A model-based control method, Model Predictive Control, has been applied in a linearised manner with robustness considerations to all these scenarios, in which a set of hard constraints must be respected. Tracking of pressure (thrust) and mixture-ratio operating points within the design envelope is achieved in simulation while respecting constraints. Robustness to variations in the parameters, which are checked to be predominant according to analyses, is also demonstrated. This framework paves the way to experimental validation via hardware-in-the-loop simulations or in test benches
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Melo, Hugo Henrique Tinoco [UNESP]. "Análise dos sprays de jatos de injetores de motor foguete utilizando um sistema de processamento digital de imagens." Universidade Estadual Paulista (UNESP), 2011. http://hdl.handle.net/11449/97043.
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A utilização de imagens digitais para extrair informações de objetos tem sido uma solução amplamente empregada em pesquisas científicas e em processos industriais. A contínua redução nos preços de equipamentos, a facilidade do uso de softwares e a simples integração com recursos de informática tem feito que muitos processos migrem para esta solução mais ágil, confiável e econômica. A indústria aeroespacial, que possui uma cadeia de produção não contínua e exige a avaliação de todos os seus componentes para obtenção de um nível de confiança elevado, encontra no emprego do processamento digital de imagens uma solução versátil e eficaz para análise das características de cada componente. Neste trabalho é apresentado um programa, desenvolvido em LabVIEW™, para medição dos sprays cônicos de jatos de injetores de motor foguete utilizando um sistema de processamento digital de imagens para sua análise. São apresentadas também as metodologias até então utilizadas para efetuar este tipo de medida. Os sprays dos jatos são desenvolvidos na saída do injetor, são exibidos visualmente durante o teste hidráulico a frio e tem influência direta no desempenho do motor foguete. A utilização desta nova ferramenta permitiu a realização desta medida de forma automática, com o fornecimento da incerteza de medição em níveis de confiança pré-estabelecido e mostrou-se ser mais exata e precisa que as metodologias anteriores
The usage of digital images to extract information from objects has been a solution widely used in scientific research and in industrial processes. The continued reduction in prices of equipment, the facility of software manipulation and the simple integration with computing resources has done many processes to migrate to this more flexible, reliable and economical solution. The aerospace industry, which has a chain of production that is not continuous and requires the evaluation of all its components to obtain a high confidence level, finds in the usage of digital image processing a versatile and effective solution for analysis of the characteristics of each component . This paper presents a program developed in LabVIEW™, to measure the rocket engine conic spray jet by using a digital image processing system for analysis. It is also presented the methodologies previously used to perform this type of measurement. The jet sprays are developed at the exit of the injector, are displayed visually during the cold hydraulic test and it has directly influences on the performance of the rocket engines. The usage of this new tool allowed us to make the measurement automatically with the supply of uncertainty together with a pre-established confidence level and it proved to be more accurate and precise than previous methodologies
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Wilson, Matthew D. "Catalytic Decomposition of Nitrous Monopropellant for Hybrid Motor Ignition." DigitalCommons@USU, 2013. https://digitalcommons.usu.edu/etd/1496.
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Nitrous oxide (N2O) is an inexpensive and readily available non-toxic rocket motor oxidizer. It is the most commonly used oxidizer for hybrid bipropellant rocket systems, and several bipropellant liquid rocket designs have also used nitrous oxide. In liquid form, N2O is highly stable, but in vapor form it has the potential to decompose exothermically, releasing up to 1865 Joules per gram of vapor as it dissociates into nitrogen and oxygen. Consequently, it has long been considered as a potential "green" replacement for existing highly toxic and dangerous monopropellants. This project investigates the feasibility of using the nitrous oxide decomposition reaction as a monopropellant energy source for igniting liquid bipropellant and hybrid rockets that already use nitrous oxide as the primary oxidizer. Because nitrous oxide is such a stable propellant, the energy barrier to dissociation is quite high; normal thermal decomposition of the vapor phase does not occur until temperatures are above 800 C. The use of a ruthenium catalyst decreases the activation energy for this reaction to allow rapid decomposition below 400 C. This research investigates the design for a prototype device that channels the energy of dissociation to ignite a laboratory scale hybrid rocket motor.
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Silva, Junior Gilberto Caetano da [UNESP]. "Método dos mínimos quadrados aplicados ao lançamento de foguetes propulsionados a ar comprimido." Universidade Estadual Paulista (UNESP), 2017. http://hdl.handle.net/11449/152165.
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Coordenação de Aperfeiçoamento de Pessoal de Nível Superior (CAPES)
Neste trabalho, apresentamos um relato de experimento realizado junto aos alunos de ensino fundamental de uma escola pública municipal e efetuamos o ajuste de curva dos dados observados por meio do método dos mínimos quadrados. Para tanto, discutimos a concepção e aplicação desse método a partir de resultados oriundos do cálculo diferencial, da álgebra linear e alguns conceitos estatísticos. Do cálculo diferencial estudamos a minimização dos erros de aproximação por meio da investigação dos pontos de mínimo da função erro. Da álgebra linear determinamos os parâmetros da função ajustada através da discussão e solução de um sistema de equações lineares resultante do conjunto de derivadas parciais nulas que estabelecem o ponto crítico da função erro. Da estatística utilizamos alguns conceitos e formulações que tratam da intensidade da relação entre as variáveis, bem como, das incertezas na variável dependente e nos parâmetros da função ajustada.
In this work, we present a report of an experiment carried out with elementary school students of a municipal public school, and we performed the curve adjustment of the observed data through the least squares method. For this, we discuss the conception and application of this method from results derived from differential calculus, linear algebra and some statistical concepts. From the differential calculation, we study the minimization of approximation errors by investigating the minimum points of the error function. From linear algebra, we determine the parameters of the adjusted function through the discussion and solution of a system of linear equations resulting from the set of null partial derivatives that establish the critical point of the error function. From statistics, we use some concepts and formulations that deal with the intensity of the relationship between variables, as well as the uncertainties in the dependent variable and the parameters of the adjusted function.
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Potier, Luc. "Large Eddy Simulation of the combustion and heat transfer in sub-critical rocket engines." Thesis, Toulouse, INPT, 2018. http://www.theses.fr/2018INPT0043/document.
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La combustion cryogénique dans les moteurs de fusée dits à propulsion liquide utilise généralement un couple d'ergols, le plus couramment composé d'hydrogène/oxygène (H2/O2). Privilégiée pour le fort pouvoir calorifique du dihydrogène, cette combustion à haute pression, induit des températures de fonctionnement très élevées et nécessite l'intégration d'un système de refroidissement. La prédiction des flux thermiques aux parois est donc un élément essentiel de la conception d'une chambre de combustion de moteur fusée. Ces flux sont le résultat d'écoulements fortement turbulents, compressibles, avec une cinétique chimique violente induisant de forts gradients d'espèces et de température. La simulation de ces phénomènes nécessite des approches spécifiques telles que la Simulation aux Grandes Echelles (SGE) qui réalise un très bon compromis entre précision et coût de calcul. Cette thèse a ainsi pour objectif la simulation par SGE des transferts de chaleur aux parois dans les chambres de combustion de moteurs fusée opérant en régime sous-critique. Le régime sous-critique implique un état liquide pour un des ergols, dont il faut traiter l'injection et l'atomisation. Dans un premier temps ce travail s'intéresse à plusieurs éléments de modélisation nécessaire pour réaliser les simulations visées. Le comportement des flammes H2/O2 est décrit par un schéma cinétique réduit et validé sur des configurations académiques. La prédictivité de ce schéma est évaluée sur une large gamme de fonctionnement dans des conditions représentatives des moteurs fusée. La simulation de l'injection de l'oxygène liquide (LOx) est un autre point critique qui nécessite de décrire l'atomisation et la phase dispersée ainsi que son couplage avec la phase gazeuse. La déstabilisation et l'atomisation primaire du jet liquide, trop complexe à simuler en SGE 3D, sont omises ici pour injecter directement un spray paramétré grâce à des corrélations empiriques. Enfin, la prédiction des flux thermiques utilise un modèle de loi de paroi spécifiquement dédiée aux écoulements à fort gradient de température. Cette loi de paroi est validée sur des configurations de canaux turbulents par comparaison avec des simulations avec résolution directe de la couche limite. La méthodologie basée sur les modèles développés est ensuite employée pour la simulation d'une chambre de combustion représentative du fonctionnement des moteurs cryogéniques. Il s'agit de la configuration CONFORTH testée sur le banc MASCOTTE (ONERA) et pour laquelle des mesures de température de paroi et de flux thermiques sont disponibles. Les résultats des SGE montrent un bon accord avec l'expérience et démontrent la capacité de la SGE à prédire les flux thermiques dans une chambre de combustion de moteur fusée. Enfin, dans un dernier chapitre ce travail s'intéresse à une méthode d'augmentation des transferts thermiques via une expérience de JAXA utilisant des parois rainurées dans la direction axiale. Par comparaison avec une chambre à parois lisses, les résultats démontrent la bonne prédiction par la SGE de l'augmentation du flux de chaleur grâce aux rainures et confirment la validité de la méthode développée pour des géométries de paroi complexesCombustion in cryogenic engines is a complex phenomenon, involving either liquid or supercritical fluids at high pressure, strong and fast oxidation chemistry, and high turbulence intensity. Due to extreme operating conditions, a particularly critical issue in rocket engine is wall heat transfer which requires efficient cooling of the combustor walls. The concern goes beyond material resistance: heat fluxes extracted through the chamber walls may be reused to reduce ergol mass or increase the power of the engine. In expander-type engine cycle, this is even more important since the heat extracted by the cooling system is used to drive the turbo-pumps that feed the chamber in fuel and oxidizer. The design of rocket combustors requires therefore an accurate prediction of wall heat flux. To understand and control the physics at play in such combustor, the Large Eddy Simulation (LES) approach is an efficient and reliable numerical tool. In this thesis work, the objective is to predict wall fluxes in a subcritical rocket engine configuration by means of LES. In such condition, ergols may be in their liquid state and it is necessary to model liquid jet atomization, dispersion and evaporation.The physics that have to be treated in such engine are: highly turbulent reactive flow, liquid jet atomization, fast and strong kinetic chemistry and finally important wall heat fluxes. This work first focuses on several modeling aspects that are needed to perform the target simulations. H2/O2 flames are driven by a very fast chemistry, modeled with a reduced mechanism validated on academic configurations for a large range of operating conditions in laminar pre- mixed and non-premixed flames. To form the spray issued from the atomization of liquid oxygen (LOx) an injection model is proposed based on empirical correlations. Finally, a wall law is employed to recover the wall fluxes without resolving directly the boundary layer. It has been specifically developed for important temperature gradients at the wall and validated on turbulent channel configurations by comparison with wall resolved LES. The above models are then applied first to the simulation of the CONFORTH sub-scale thrust chamber. This configuration studied on the MASCOTTE test facility (ONERA) has been measured in terms of wall temperature and heat flux. The LES shows a good agreement compared to experiment, which demonstrates the capability of LES to predict heat fluxes in rocket combustion chambers. Finally, the JAXA experiment conducted at JAXA/Kakuda space center to observe heat transfer enhancement brought by longitudinal ribs along the chamber inner walls is also simulated with the same methodology. Temperature and wall fluxes measured with smooth walls and ribbed walls are well recovered by LES. This confirms that the LES methodology proposed in this work is able to handle wall fluxes in complex geometries for rocket operating conditions
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Messineo, Jérôme. "Modélisation des instabilités hydrodynamiques dans les moteurs-fusées hybrides." Thesis, Toulouse, ISAE, 2016. http://www.theses.fr/2016ESAE0025/document.
Full textAbstract:
Les moteurs-fusées hybrides combinent les technologies des deux autres catégories de moteurs à propulsionchimique, et associent un combustible et un oxydant stockés respectivement sous phase solide et liquide.Cette architecture offre un certain nombre d’avantages, comme par exemple des coûts plus faibleset une architecture simplifiée par rapport à la propulsion bi-liquide; la possibilité de réaliser de multiplesextinctions et ré-allumages et une bonne impulsion spécifique théorique par rapport à la propulsion solide,et enfin une sécurité de mise en œuvre accrue et un impact environnemental faible vis-à-vis de ces deuxautres modes de propulsion. Comme toutes les chambres de combustion, celles des moteurs hybrides peuvent subir des oscillations de pression sous certaines conditions de fonctionnement. Ces instabilités se traduisent par des fluctuationsde poussée qui peuvent dégrader la structure d’un lanceur ou d’un satellite. Des phénomènes diverspeuvent être à l’origine des fluctuations de pression observées dans les moteurs hybrides.L’objectif de la thèse est de proposer une modélisation des instabilités d’origine hydrodynamique quiapparaissent dans les moteurs hybrides. Une exploitation nouvelle de la base de données disponible àl’ONERA a servi de support pour la modélisation, ainsi que des simulations numériques instationnaires 2Det 3D réalisées à l’aide du code CFD CEDRE. Les instabilités sont provoquées par la formation périodiquede structures tourbillonnaires dans la chambre de combustion, qui génèrent des fluctuations de pressionlors de leur passage dans le col de la tuyère. L’originalité du modèle, basé sur la théorie classique degénération tourbillonnaire dans une cavité, consiste à prendre en compte les variations géométriques dela chambre de combustion au cours des tirs. Ces variations ont un effet sur la vitesse de l’écoulement, surla zone de recirculation dans la post-chambre, ainsi que sur les tourbillons eux-mêmes. Enfin, plusieursnouveaux essais du moteur hybride HYCOM ont été effectués et confrontés au modèle développé dans lecadre de la thèseHybrid rocket motors combine solid and bi-liquid chemical propulsion technologies and associate asolid fuel and a liquid oxidizer in its classical configuration. This architecture offers several advantagesover liquid propulsion such as lower costs and a simplified architecture. The possibility of performingmultiple extinctions and re-ignitions and a good theoretical specific impulse is also an improvement inregard to solid propulsion. Hybrid engines also have improved safety and a lower environmental impactthan other chemical propulsion systems. As in all combustion chambers, hybrid engines suffer from pressure oscillations under specific operating conditions. These instabilities provoke thrust fluctuations that can damage the launcher and payloads.Various phenomena can induce the pressure oscillations observed in hybrid rocket engines.The objective of this thesis is to propose a model of hydrodynamics instabilities that appear in hybridengines. A new exploitation of the database available at ONERA, and unsteady 2D and 3D numericalsimulations were used for the modeling. The instabilities are provoked by the periodic formation ofvortices in the combustion chamber that generate pressure fluctuations when passing through the nozzlethroat. The originality of the model, which is based on the classical theory of vortices generation ina cavity, consists in taking into account the geometrical variations of the combustion chamber duringoperation. These variations have an effect on the flow velocity, on the recirculation area in the postchamberand on the vortices. Finally, several new firing tests of the hybrid engine HYCOM have beenperformed and compared to the model developed in this thesis
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Josselyn, Scott B. "Optimization of low thrust trajectories with terminal aerocapture." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Jun%5FJosselyn.pdf.
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Thesis (Aeronautical and Astronautical Engineer)--Naval Postgraduate School, June 2003.Thesis advisor(s): I. Michael Ross, Steve Matousek. Includes bibliographical references (p. 149-150). Also available online.
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Morham, Brett G. "Numerical Examination of Flow Field Characteristics and Fabri Choking of 2D Supersonic Ejectors." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/340.
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An automated computer simulation of the two-dimensional planar Cal Poly Supersonic Ejector test rig is developed. The purpose of the simulation is to identify the operating conditions which produce the saturated, Fabri choke and Fabri block aerodynamic flow patterns. The effect of primary to secondary stagnation pressure ratio on the efficiency of the ejector operation is measured using the entrainment ratio which is the secondary to primary mass flow ratio.
The primary flow of the ejector is supersonic and the secondary (entrained) stream enters the ejector at various velocities at or below Mach 1. The primary and secondary streams are both composed of air. The primary plume boundary and properties are solved using the Method of Characteristics. The properties within the secondary stream are found using isentropic relations along with stagnation conditions and the shape of the primary plume. The solutions of the primary and secondary streams iterate on a pressure distribution of the secondary stream until a converged solution is attained. Viscous forces and thermo-chemical reactions are not considered.
For the given geometry the saturated flow pattern is found to occur below stagnation pressure ratios of 74. The secondary flow of the ejector becomes blocked by the primary plume above pressure ratios of 230. The Fabri choke case exists between pressure ratios of 74 and 230, achieving optimal operation at the transition from saturated to Fabri choked flow, near the pressure ratio of 74. The case of optimal expansion yields an entrainment ratio of 0.17. The entrainment ratio results of the Cal Poly Supersonic Ejector simulation have an average error of 3.67% relative to experimental data. The accuracy of this inviscid simulation suggests ejector operation in this regime is governed by pressure gradient rather than viscous effects.
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Deans, Matthew Charles. "The Simulation, Development, and Testing of a Staged Catalytic Microtube Ignition System." Case Western Reserve University School of Graduate Studies / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=case1343318716.
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Sarotte, Camille. "Improvement of monitoring and reconfiguration processes for liquid propellant rocket engine." Thesis, Université Paris-Saclay (ComUE), 2019. http://www.theses.fr/2019SACLS348/document.
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La surveillance et l'amélioration des modes de fonctionnement des systèmes propulsifs des lanceurs représentent des défis majeurs de l'industrie aérospatiale. En effet, une défaillance ou un dysfonctionnement du système propulsif peut avoir un impact significatif pour les clients institutionnels ou privés et entraîner des catastrophes environnementales ou humaines. Des systèmes de gestion de la santé (HMS) pour les moteurs fusée à ergols liquides (LPREs), ont été mis au point pour tenir compte des défis actuels en abordant les questions de sureté et de fiabilité. Leur objectif initial est de détecter les pannes ou dysfonctionnements, de les localiser et de prendre une décision à l’aide de Redlines et de systèmes experts. Cependant, ces méthodes peuvent induire de fausses alarmes ou des non-détections de pannes pouvant être critiques pour la sécurité et la fiabilité des opérations. Ainsi, les travaux actuels visent à éliminer certaines pannes critiques, mais aussi diminuer les arrêts intempestifs. Les données disponibles étant limitées, des méthodes à base de modèles sont essentiellement utilisées. La première tâche consiste à détecter les défaillances de composants et/ou d'instruments à l'aide de méthodes de détection et de localisation de fautes (FDI). Si la faute est considérée comme mineure, des actions de « non-arrêt » sont définies pour maintenir les performances de l'ensemble du système à un niveau proche de celles souhaitées et préserver les conditions de stabilité. Il est donc nécessaire d’effectuer une reconfiguration robuste (incertitudes, perturbations inconnues) du moteur. Les saturations en entrée doivent également être prises en compte dans la conception de la loi de commande, les signaux de commande étant limités en raison des caractéristiques ou performances des actionneurs physiques. Les trois objectifs de cette thèse sont donc : la modélisation des différents sous-systèmes principaux d’un LPRE, le développement d’algorithmes de FDI sur la base des modèles établis et la définition d’un système de reconfiguration du moteur en temps réel pour compenser certains types de pannes. Le système de FDI et Reconfiguration (FDIR) développé sur la base de ces trois objectifs a ensuite été validé à l’aide de simulations avec CARINS (CNES) et du banc d’essai MASCOTTE (CNES/ONERA)Monitoring and improving the operating modes of launcher propulsion systems are major challenges in the aerospace industry. A failure or malfunction of the propulsion system can have a significant impact for institutional or private customers and results in environmental or human catastrophes. Health Management Systems (HMS) for liquid propellant rocket engines (LPREs), have been developed to take into account the current challenges by addressing safety and reliability issues. Their objective was initially to detect failures or malfunctions, isolate them and take a decision using Redlines and Expert Systems. However, those methods can induce false alarms or undetected failures that can be critical for the operation safety and reliability. Hence, current works aim at eliminating some catastrophic failures but also to mitigate benign shutdowns to non-shutdown actions. Since databases are not always sufficient to use efficiently data-based analysis methods, model-based methods are essentially used. The first task is to detect component and / or instrument failures with Fault Detection and Isolation (FDI) approaches. If the failure is minor, non-shutdown actions must be defined to maintain the overall system current performances close to the desirable ones and preserve stability conditions. For this reason, it is required to perform a robust (uncertainties, unknown disturbances) reconfiguration of the engine. Input saturation should also be considered in the control law design since unlimited control signals are not available due to physical actuators characteristics or performances. The three objectives of this thesis are therefore: the modeling of the different main subsystems of a LPRE, the development of FDI algorithms from the previously developed models and the definition of a real-time engine reconfiguration system to compensate for certain types of failures. The developed FDI and Reconfiguration (FDIR) scheme based on those three objectives has then been validated with the help of simulations with CARINS (CNES) and the MASCOTTE test bench (CNES/ONERA)
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Garby, Romain. "Simulations of flame stabilization and stability in high-pressure propulsion systems." Phd thesis, Toulouse, INPT, 2013. http://oatao.univ-toulouse.fr/9706/1/garby.pdf.
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Rust, Jack W. "Fuel optimal low thrust trajectories for an asteroid sample return mission." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2005. http://library.nps.navy.mil/uhtbin/hyperion/05Mar%5FRust.pdf.
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Melo, Hugo Henrique Tinoco. "Análise dos sprays de jatos de injetores de motor foguete utilizando um sistema de processamento digital de imagens /." Guaratinguetá : [s.n.], 2011. http://hdl.handle.net/11449/97043.
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Orientador: Fernando de Azevedo SilvaBanca: João Zangrandi Filho
Banca: Silvana Aparecida Barbosa
Resumo: A utilização de imagens digitais para extrair informações de objetos tem sido uma solução amplamente empregada em pesquisas científicas e em processos industriais. A contínua redução nos preços de equipamentos, a facilidade do uso de softwares e a simples integração com recursos de informática tem feito que muitos processos migrem para esta solução mais ágil, confiável e econômica. A indústria aeroespacial, que possui uma cadeia de produção não contínua e exige a avaliação de todos os seus componentes para obtenção de um nível de confiança elevado, encontra no emprego do processamento digital de imagens uma solução versátil e eficaz para análise das características de cada componente. Neste trabalho é apresentado um programa, desenvolvido em LabVIEW™, para medição dos sprays cônicos de jatos de injetores de motor foguete utilizando um sistema de processamento digital de imagens para sua análise. São apresentadas também as metodologias até então utilizadas para efetuar este tipo de medida. Os sprays dos jatos são desenvolvidos na saída do injetor, são exibidos visualmente durante o teste hidráulico a frio e tem influência direta no desempenho do motor foguete. A utilização desta nova ferramenta permitiu a realização desta medida de forma automática, com o fornecimento da incerteza de medição em níveis de confiança pré-estabelecido e mostrou-se ser mais exata e precisa que as metodologias anteriores
Abstract: The usage of digital images to extract information from objects has been a solution widely used in scientific research and in industrial processes. The continued reduction in prices of equipment, the facility of software manipulation and the simple integration with computing resources has done many processes to migrate to this more flexible, reliable and economical solution. The aerospace industry, which has a chain of production that is not continuous and requires the evaluation of all its components to obtain a high confidence level, finds in the usage of digital image processing a versatile and effective solution for analysis of the characteristics of each component . This paper presents a program developed in LabVIEW™, to measure the rocket engine conic spray jet by using a digital image processing system for analysis. It is also presented the methodologies previously used to perform this type of measurement. The jet sprays are developed at the exit of the injector, are displayed visually during the cold hydraulic test and it has directly influences on the performance of the rocket engines. The usage of this new tool allowed us to make the measurement automatically with the supply of uncertainty together with a pre-established confidence level and it proved to be more accurate and precise than previous methodologies
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