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Relevant bibliographies by topics / Rocket propulsion. Rockets

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Academic literature on the topic 'Rocket propulsion. Rockets'

Author: Grafiati

Published: 4 June 2021

Last updated: 1 February 2022

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Journal articles on the topic "Rocket propulsion. Rockets"

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Wang, Yu Fei, Gong Chen, and Li Lin Han. "THE Comprehensive Survey for the Numerical Simulation of the 4th Generation Rocket Ejection Seat Thrust Vector Control System." Applied Mechanics and Materials 551 (May 2014): 523–29. http://dx.doi.org/10.4028/www.scientific.net/amm.551.523.

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The objective of this work was to develop and adapt existing computational methodologies for 3D comprehensive analysis of ejection seat aerodynamics, including rocket plume effects.Various methods were investigated for prescribing boundary and initial conditions for the seat rockets. The selected method utilizes a model that prescribes 3D nozzle exit boundary profiles extracted from detailed rocket nozzle calculations. Also a multi-domain gridding method that allows for many-to-one interface meshing was developed and tested for efficient and accurate rocket plume resolution within the 3D ejection seat computational environment. Validations were performed for Aerojet’s Pintle Escape Propulsion system (PEPS) rockets. Three dimensional ejection seat calculations with rocket power were also made to demonstrate the feasibility of the approach and the potential use of the model.

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Haw, Stephen G. "Cathayan Arrows and Meteors: The Origins of Chinese Rocketry." Journal of Chinese Military History 2, no. 1 (2013): 28–42. http://dx.doi.org/10.1163/22127453-12341243.

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Abstract Although it is now generally accepted that the rocket was invented in China, there is little agreement about exactly when this occurred. Various conflicting claims have been made, usually on the basis of dubious evidence. This article examines these claims and rejects all of them. It also discounts the hypothesis that the rocket developed from a kind of firework called a “fire-rat.” Instead, it suggests that the rocket very probably developed from Chinese fire-arrows, which carried charges of incendiary gunpowder. These became rocket-assisted arrows, fired from bows or, very often, arbalests. They achieved great ranges, of the order of two miles. They were used by the Mongols during their conquests in the mid-thirteenth century and may have formed part of the weaponry of the Chinese Song dynasty at an earlier period. The earliest reference to what were very probably true rockets, launched by their own propulsion, dates from 1272. It occurs in an account of events that took place during the siege of Xiangyang by Mongol forces. These rockets were used for signaling. There is no evidence that self-launching rockets were utilized as weapons in the mid-thirteenth century. Rockets as weapons are unlikely to have been developed any earlier than the late 1200s.

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Bolivar, Nelson Enrique, and Ivaylo T. Vasilev. "Non-Combustion ⁴He Powered Propulsion." European Journal of Engineering and Technology Research 6, no. 2 (February 16, 2021): 101–6. http://dx.doi.org/10.24018/ejers.2021.6.2.2283.

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One of the biggest hurdles nowadays rocket propulsion is the large use of fuel. The amount of fuel and the burning efficiency defines how long the rocket engine can work which intimately limits the range and the load capacity of the rockets and spaceships. This according to the Newton third law is unavoidable - in order to move forward you need to leave something behind. There have been several attempts in the past to create an engine which doesn't use fuel in the common sense, like the M drive, but so far all of them were unsuccessful. In this article we attempt to explore a novel principle, a recycling cycle of fuel, by optimizing parametrically a system that uses ⁴He phase transition.

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Freiherr, Greg. "The Little Rocket Engine That Could." Mechanical Engineering 138, no. 08 (August 1, 2016): 32–37. http://dx.doi.org/10.1115/1.2016-aug-1.

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This paper highlights advancements of a research team in the field of miniature spacecrafts and development and advantages of CubeStats. CubeSats are space age hitchhikers, that is, miniature spacecraft that fly into orbit aboard rockets whose primary payloads are full-size satellites. Paulo Lozano and his team at MIT’s Space Propulsion Lab have developed a unique kind of rocket engine for these microsatellites. The trick to building a successful ion electrospray propulsion system is to increase thrust density by jamming together as many emitters as possible. Electrospray engines also differ greatly from another form of ion propulsion, plasma ion, which also eschews chemical combustion for the efficiency of the electron. A big advantage, when asking for permission to hitchhike a ride into orbit, is that ion electrospray propulsion engines cannot explode and destroy a rocket’s primary payload. A second advantage of ion engines is their modularity and, consequently, scalability.

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Apel, Uwe, Alexander Baumann, Christian Dierken, and Thilo Kunath. "AQUASONIC – A Sounding Rocket Based on Hybrid Propulsion." Applied Mechanics and Materials 831 (April 2016): 3–13. http://dx.doi.org/10.4028/www.scientific.net/amm.831.3.

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The AQUASONIC project is aimed to develop a sounding rocket including a hybrid propulsion system based on the propellant combination nitrous oxide and polyethylene. It takes place in the frame of the STERN (Student Experimental Rockets) programme founded by the German Space Agency (DLR) in order to promote students in the area of launch vehicles. Main element of the project is the AQUASONIC rocket, which shall reach a flight altitude of 5-6 km and a velocity of MACH 1. All major activities like design, manufacturing, verification and, finally, the launch campaign will be performed by students. The rocket shall be launched at Esrange Space Centre (Sweden) in 2016. Thus, students are able to apply their skills and knowledge to a real project like it is conducted by the space industry or research organisations.

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Santos, L. M. C., L. A. R. Almeida, A. M. Fraga, and C. A. G. Veras. "EXPERIMENTAL INVESTIGATION OF A PARAFFIN BASED HYBRID ROCKET." Revista de Engenharia Térmica 5, no. 1 (July 31, 2006): 08. http://dx.doi.org/10.5380/reterm.v5i1.61658.

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Hybrid rockets are known to be simpler, safer, environmentally friend, and, more importantly, cheaper than most of the technologies for propulsion devices used today. Hybrid rockets can be applied as the propulsion system in satellites launch vehicles, micro-satellites and tactical missiles. This paper deals with combustion of ultra-high molecular weight polyethylene (UHMWPE) and paraffin as the solid fuels burning with gaseous oxygen (GOX) as well as N O as the oxidizer in lab scale hybrid rocket motors. A test 2 stand was built to carry out the experiments. The main objectives were to investigate the ignition of the solid fuels, burning performance and regression rates for different operating conditions. With paraffin-based fuel the hybrid motor had the regression rate enhanced two to three folds compared to the UHMWPE, as reported in the literature. The overall performance of the motor, with paraffin as the fuel, is comparable to other technologies. Paraffin-based hybrid rockets can, then, be a safer and cheaper alternative to satellite launch vehicles for the Brazilian space program.

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Okninski, Adam, Pawel Surmacz, Bartosz Bartkowiak, Tobiasz Mayer, Kamil Sobczak, Michal Pakosz, Damian Kaniewski, Jan Matyszewski, Grzegorz Rarata, and Piotr Wolanski. "Development of Green Storable Hybrid Rocket Propulsion Technology Using 98% Hydrogen Peroxide as Oxidizer." Aerospace 8, no. 9 (August 24, 2021): 234. http://dx.doi.org/10.3390/aerospace8090234.

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This paper presents the development of indigenous hybrid rocket technology, using 98% hydrogen peroxide as an oxidizer. Consecutive steps are presented, which started with interest in hydrogen peroxide and the development of technology to obtain High Test Peroxide, finally allowing concentrations of up to 99.99% to be obtained in-house. Hydrogen peroxide of 98% concentration (mass-wise) was selected as the workhorse for further space propulsion and space transportation developments. Over the course nearly 10 years of the technology’s evolution, the Lukasiewicz Research Network—Institute of Aviation completed hundreds of subscale hybrid rocket motor and component tests. In 2017, the Institute presented the first vehicle in the world to have demonstrated in-flight utilization for 98% hydrogen peroxide. This was achieved by the ILR-33 AMBER suborbital rocket, which utilizes a hybrid rocket propulsion as the main stage. Since then, three successful consecutive flights of the vehicle have been performed, and flights to the Von Karman Line are planned. The hybrid rocket technology developments are described. Advances in hybrid fuel technology are shown, including the testing of fuel grains. Theoretical studies and sizing of hybrid propulsion systems for spacecraft, sounding rockets and small launch vehicles have been performed, and planned further developments are discussed.

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Heeg, Francesca, Lukas Kilzer, Robin Seitz, and Enrico Stoll. "Design and Test of a Student Hybrid Rocket Engine with an External Carbon Fiber Composite Structure." Aerospace 7, no. 5 (May 13, 2020): 57. http://dx.doi.org/10.3390/aerospace7050057.

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The development of hybrid rockets offers excellent opportunities for the practical education of students at universities due to the high safety and relatively low complexity of the rocket propulsion system. During the German educational program Studentische Experimental-Raketen (STERN), students of the Technische Universität Braunschweig obtain the possibility to design and launch a sounding rocket with a hybrid engine. The design of the engine HYDRA 4X (HYbridDemonstrations-RaketenAntrieb) is presented, and the results of the first engine tests are discussed. The results for measured regression rates are compared to the results from the literature. Furthermore, the impact of the lightweight casing material carbon fiber-reinforced plastic (CFRP) on the hybrid engine mass and flight apogee altitude is examined for rockets with different total impulse classes (10 to 50 kNs). It is shown that the benefit of a lightweight casing material on engine mass decreases with an increasing total impulse. However, a higher gain on apogee altitude, especially for bigger rockets with a comparable high total impulse, is shown.

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Zaraini, Fairul Azmin, Tengku Farah Wahida Ku Chik, Nor Hafizah Abdullah, and Ahmad Ammar. "Suggestions for a Roadmap towards Becoming a Launch Capable Nation." Applied Mechanics and Materials 225 (November 2012): 561–65. http://dx.doi.org/10.4028/www.scientific.net/amm.225.561.

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Propulsion Technology Program under the National Space Agency (ANGKASA) was commenced in 2009 with an ambition to launch satellites into Low Earth Orbit (LEO) using its own independent launch vehicle. Four members of the Space Application and Technology Development (SATD) with various backgrounds have been entrusted to draft roadmap for National Satellite Launcher and at the same time conducting Research and Development (R&D) related to rocketry. The first program was solid rocket development between ANGKASA and Universiti Teknologi Malaysia (UTM) through budget allocated in Rancangan Malaysia ke-9 (RMK-9). The rockets developed in this project have been successfully launched at eastern coast of peninsula Malaysia in 2010. This achievement needs proper and effective continuation towards enabling Malaysia to be a launch capable nation. Therefore, this paper investigates rocket development programs and activities ran by various countries which could be adopted into national programs in order to spur participation in rocket science and space industries, hence materialise completion of Malaysian own launch vehicle in timely manner. Moreover, this paper will look over obstacles and potencies of rocket development with current Malaysian environment.

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Asraff, A. K., S. Sheela, Krishnajith Jayamani, S. Sarath Chandran Nair, and R. Muthukumar. "Material Characterisation and Constitutive Modelling of a Copper Alloy and Stainless Steel at Cryogenic and Elevated Temperatures." Materials Science Forum 830-831 (September 2015): 242–45. http://dx.doi.org/10.4028/www.scientific.net/msf.830-831.242.

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High performance rockets are developed using cryogenic technology. High thrust cryogenic rocket engines operating at elevated temperatures and pressures are the backbone of such rockets. The thrust chamber of such engines, which produce the thrust for the propulsion of the rocket, can be considered as structural elements. Often double walled construction is employed for these chambers for better cooling and enhanced performance. The thrust chamber investigated here has its hot inner wall fabricated out of a high conductivity high ductility copper alloy and outer wall made of a ductile stainless steel. The engine is indigenously designed and developed by ISRO and is undergoing hot tests. Inner wall is subjected to high thermal and pressure loads during operation of engine due to which it will be in the plastic regime. Evaluation of tensile properties of the copper alloy and stainless steel up to fracture, at cryogenic, ambient and elevated temperatures in parent metal and welded forms is of paramount importance for its constitutive modelling and thermo structural analysis of the thrust chamber.

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Dissertations / Theses on the topic "Rocket propulsion. Rockets"

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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.

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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).

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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.

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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.

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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.

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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.

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Books on the topic "Rocket propulsion. Rockets"

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Oscar, Biblarz, ed. Rocket propulsion elements. 7th ed. New York: John Wiley & Sons, 2001.

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Rocket propulsion elements: An introduction to the engineering of rockets. 5th ed. New York: Wiley, 1986.

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Rocket propulsion elements: An introduction to the engineering of rockets. 6th ed. New York: Wiley, 1992.

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Timnat, Y. M. Advanced chemical rocket propulsion. London: Academic Press, 1987.

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Timnat, Y. M. Advanced chemical rocket propulsion. London: Academic Press, 1987.

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International Electric Propulsion Conference (24th 1995 Moscow, Russia). Proceedings of the 24th International Electric Propulsion Conference: IEPC. [S.l: s.n., 1995.

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Dawson, Virginia P. Rocket propulsion research at Lewis Research Center. [Washington, DC: National Aeronautics and Space Administration, 1992.

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Dawson, Virginia P. Rocket propulsion research at Lewis Research Center. [Washington, DC: National Aeronautics and Space Administration, 1992.

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Rocket and spacecraft propulsion: Principles, practice and new developments. 2nd ed. Berlin: Springer, 2005.

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Elements of propulsion: Gas turbines and rockets. Reston, Va: American Institute of Aeronautics and Astronautics, 2006.

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Book chapters on the topic "Rocket propulsion. Rockets"

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Karabeyoğlu, Arif. "Performance Additives for Hybrid Rockets." In Chemical Rocket Propulsion, 139–63. Cham: Springer International Publishing, 2016. http://dx.doi.org/10.1007/978-3-319-27748-6_5.

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El-Sayed, Ahmed F. "Rocket Propulsion." In Fundamentals of Aircraft and Rocket Propulsion, 907–91. London: Springer London, 2016. http://dx.doi.org/10.1007/978-1-4471-6796-9_11.

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Nixon, John. "Electrical Rocket Propulsion." In Modern English for Aeronautics and Space Technology, 102–12. München: Carl Hanser Verlag GmbH & Co. KG, 2011. http://dx.doi.org/10.3139/9783446428348.009.

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Mishra, D. P. "Rocket Nozzle." In Fundamentals of Rocket Propulsion, 91–127. Boca Raton: CRC Press, 2017.: CRC Press, 2017. http://dx.doi.org/10.1201/9781315175997-4.

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Denny, Mark, and Alan McFadzean. "Rocket Propulsion and Guidance." In Rocket Science, 143–70. Cham: Springer International Publishing, 2019. http://dx.doi.org/10.1007/978-3-030-28080-2_5.

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Mishra, D. P. "Chemical Rocket Propellants." In Fundamentals of Rocket Propulsion, 161–94. Boca Raton: CRC Press, 2017.: CRC Press, 2017. http://dx.doi.org/10.1201/9781315175997-6.

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Mishra, D. P. "Nonchemical Rocket Engine." In Fundamentals of Rocket Propulsion, 397–438. Boca Raton: CRC Press, 2017.: CRC Press, 2017. http://dx.doi.org/10.1201/9781315175997-11.

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Sjovold, Arve R., and Damon C. Morrison. "Rocket Propulsion Cost Modeling." In Cost Analysis Applications of Economics and Operations Research, 226–58. New York, NY: Springer US, 1989. http://dx.doi.org/10.1007/978-1-4684-6384-2_14.

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Mishra, D. P. "Elements of Rocket Propulsion." In Fundamentals of Rocket Propulsion, 69–90. Boca Raton: CRC Press, 2017.: CRC Press, 2017. http://dx.doi.org/10.1201/9781315175997-3.

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Hori, Keiichi. "Lessons Learned in the Thruster Tests of HAN." In Chemical Rocket Propulsion, 801–18. Cham: Springer International Publishing, 2016. http://dx.doi.org/10.1007/978-3-319-27748-6_33.

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Conference papers on the topic "Rocket propulsion. Rockets"

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Naumann, Karl W., Matthias Berndl, Ludwig Eineder, Raphael Esterl, Guenter Fechler, Andreas Hacker, Tobias Meyer, et al. "A First Stage Solid Propellant Rocket Motor for Sounding Rockets." In AIAA Propulsion and Energy 2021 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2021. http://dx.doi.org/10.2514/6.2021-3692.

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Naumann, Karl W., Matthias Berndl, Ludwig Eineder, Raphael Esterl, Guenter Fechler, Andreas Hacker, Tobias Meyer, et al. "Correction: A First Stage Solid Propellant Rocket Motor for Sounding Rockets." In AIAA Propulsion and Energy 2021 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2021. http://dx.doi.org/10.2514/6.2021-3692.c1.

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Casalino, Lorenzo, and Dario Pastrone. "A Parametric Analysis of Hybrid Rocket Motors for Sounding Rockets." In 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2008. http://dx.doi.org/10.2514/6.2008-4544.

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Takahashi, Koji. "Contemporary Technology and Application of MEMS Rocket." In CANEUS 2006: MNT for Aerospace Applications. ASMEDC, 2006. http://dx.doi.org/10.1115/caneus2006-11031.

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MEMS rocket array is one of the most promising propulsion systems for micro satellite. Its recent development researches have reached close to practical devices. This paper reviews the performances and maturities of recently-build MEMS rockets. Four MEMS rocket projects are treated; LAAS-CNRS, France, Tohoku University, Japan, National University of Singapore, and Kyushu University, Japan. Their design, size, propellant, etc. are explained. The smallest MEMS rocket of Kyushu University using DDNP propellant and PDMS micro tank is also introduced.

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Nguyen, Bao, Khulood Faruqui, Luis R. Robles, Johnny Ho, Geoffrey Wagner, Jeremy Surmi, Ashley Carter, et al. "Overview of Current Hybrid Propulsion Research and Development." In ASME 2017 International Mechanical Engineering Congress and Exposition. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/imece2017-72429.

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As dwindling federal funding continues to constrict the national space program, private entities have carried the torch of innovation in the aerospace industry. While the concept of hybrid rocket engines, systems where solid fuels and fluid oxidizers are used for combustion, was conceived during the mid-20th century, the aerospace industry only recently has substantially increased research and development of these engines. According to the literature, hybrids are safer and cheaper than their liquid counterparts due to the utilization of solid fuel and generally provide greater values of specific impulse, density specific impulse, and fuel energy density than traditional solid-fuel engines. This paper provides an overview of the design principles used to develop hybrid engines and discusses limitations currently faced by industry. Furthermore, development of hybrid engines to allow for both scalability and reusability are explored as private aerospace companies continue to demonstrate that reusable rockets are the future of rocket technology. With applications catering from low-payload cargo delivery to the increasing interest in space tourism and exploration, hybrid engines can provide a safer, less expensive solution than traditional and well-established engine selections. Suggestions for future design opportunities and methods are proposed and discussed to make the concept of hybrid engines a viable innovation in the future of rocketry.

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Steppert, Michael, and Philipp Epple. "Numerical Investigation of the Drag of Rockets at Subsonic, Transonic and Supersonic Speeds." In ASME 2017 International Mechanical Engineering Congress and Exposition. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/imece2017-71872.

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To compute the trajectory of a rocket the knowledge of the external aerodynamics and the resulting forces is essential. The drag coefficient is an important parameter in the computation of rocket trajectory in case of vertical ascent. For compressible flows, at subsonic and supersonic velocities, the drag coefficient is a function of the Reynolds number and of the Mach number. Both dimensionless numbers depend on the temperature, however, not in the same way. For that reason, the Mach number dependency and the Reynolds number dependency are different. Since in the atmosphere pressure and temperature are functions of the height, the drag of rockets is also dependent on the height of the rocket in the atmosphere. In this work, the dependence of the drag on the shape of the rocket is investigated at different heights and velocities, i.e. Reynolds and Mach numbers. For this purpose, the historical rocket from Johannes Winkler, the first liquid propulsion rocket in Europe, and a modern rocket geometry were chosen and compared. These rockets were investigated numerically with the commercial Navier–Stokes solver STAR-CCM+. A detailed analysis of the drag coefficient, split into friction, pressure and wave drag was performed at these heights, Mach numbers and Reynolds numbers for the different aerodynamic shapes and rockets. In particular on the transonic and supersonic range the shock wave system leading to the wave drag was analysed in detail. Graphs of the corresponding friction, pressure, wave and total drag coefficients as a function of the Reynolds and as a function of the Mach numbers at different heights and a detailed analysis of these results are shown for the two rockets, the historical and a modern one.

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Ojeda, Carlos, Kenton T. Prescott, and Tekilanand Persaud. "Production and Manufacture of Low-Cost Liquid Rocket Engines for Sounding Rockets." In 53rd AIAA/SAE/ASEE Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2017. http://dx.doi.org/10.2514/6.2017-4841.

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Wang, Jiyuan, Longqiu Li, Xiaocong Chang, Tianlong Li, Wenping Song, and Guangyu Zhang. "The Effect of Geometry on the Velocity and Drag Force of Catalytic Micro/Nano-Rockets." In ASME 2015 International Design Engineering Technical Conferences and Computers and Information in Engineering Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/detc2015-46881.

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The hydrodynamic behavior of synthetic self-propelled catalytic micro/nano-rocket moving in low Reynolds number flow is studied theoretically. The inclination angle of the bubble departed from the micro/nano-rocket is related to the radius of the micro/nano-rocket. A unified formula of the drag force for cylindrical, cone-frustum and double truncated cone shapes micro/nano-rocket have been derived. The effect of geometric shapes on the velocity and the drag force is identified by comparing with three circular cross-sectional types of micro/nano-rocket. The average velocity is found to be strongly dependent on semi-cone angle, length, radius of the micro/nano-rocket, the H2O2 concentration and the drag force. This model provides a proficient explanation for propulsion mechanism of a catalytic micro/nano-rocket. This work can be used to optimize catalytic micro/nano-rockets design, which may have potential applications in biomedical and environmental engineering.

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TKACHENKO, YURIJ, and CHARLES LIMERICK. "Powerful liquid rocket engine (LRE) created by NPO Energomash for upto date space rockets." In 29th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1993. http://dx.doi.org/10.2514/6.1993-1957.

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Kobald, M., C. Schmierer, U. Fischer, K. Tomilin, A. Petrarolo, and M. Rehberger. "The HyEnD stern hybrid sounding rocket project." In Progress in Propulsion Physics – Volume 11. Les Ulis, France: EDP Sciences, 2019. http://dx.doi.org/10.1051/eucass/201911025.

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The student team Hybrid Engine Development (HyEnD) of the University of Stuttgart is taking part with the Institute of Space Systems (IRS) in the DLR educational program STERN (Studentische Experimentalraketen). This program supports students at German universities to design, build, and launch an experimental rocket within a 3-year project time frame. HyEnD is developing a hybrid rocket called HEROS (Hybrid Experimental Rocket Stuttgart) with a design thrust of 10 kN, a total impulse of over 100 kN·s, and an expected liftoff weight up to 175 kg. HEROS is planned to be launched in October 2015 from Esrange in Sweden to an expected flight altitude of 40 to 50 km. The current altitude record for amateur rockets in Europe is at approximately 21 km. The propulsion system of HEROS is called HyRES (Hybrid Rocket Engine Stuttgart) and uses a paraffin-based solid fuel and nitrous oxide (N2O) as a liquid oxidizer. The development and the test campaign of HyRES is described in detail. The main goals of the test campaign are to achieve a combustion efficiency higher than 90% and provide stable operation with low combustion chamber pressure fluctuations. The successful design and testing of the HyRES engine was enabled by the evaluation and characterization of a small-scale demonstrator engine. The 500-newton hybrid rocket engine, called MIRAS (MIcro RAkete Stuttgart), has also been developed in the course of the STERN project as a technology demonstrator. During this test campaign, a ballistic characterization of paraffin-based hybrid rocket fuels with different additives in combination with N2O and a performance evaluation were carried out. A wide range of operating conditions, fuel compositions, injector geometries, and engine configurations were evaluated with this engine. Effects of different injector geometries and postcombustion chamber designs on the engine performance were analyzed. Additionally, the appearance of combustion instabilities under certain conditions, their effects, and possible mitigation techniques were also investigated. Concluding, the development and construction of an advanced, lightweight hybrid sounding rocket for the given requirements and budget within the DLR STERN program are described herein. The most important parts include a high thrust hybrid rocket engine, the development of a light weight oxidizer tank, pyrotechnical valves, carbon fiber rocket structure, recovery systems, and onboard electronics.

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Reports on the topic "Rocket propulsion. Rockets"

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AIR FORCE RESEARCH LAB EDWARDS AFB CA. Propulsion and Energy: Solid Rockets. Fort Belvoir, VA: Defense Technical Information Center, August 2001. http://dx.doi.org/10.21236/ada404865.

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Lawrence, Timothy J. Nuclear Thermal Rocket Propulsion Systems. Fort Belvoir, VA: Defense Technical Information Center, March 2005. http://dx.doi.org/10.21236/ada430931.

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Watson, C. W. Nuclear rockets: High-performance propulsion for Mars. Office of Scientific and Technical Information (OSTI), May 1994. http://dx.doi.org/10.2172/10160494.

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Umholtz, Philip D. The History of Solid Rocket Propulsion and Aerojet. Fort Belvoir, VA: Defense Technical Information Center, April 1999. http://dx.doi.org/10.21236/ada406104.

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DeGeorge, Drew, and Scott Fletcher. The Integrated High Payoff Rocket Propulsion Technology Program and Tactical Missile Propulsion Status. Fort Belvoir, VA: Defense Technical Information Center, September 2002. http://dx.doi.org/10.21236/ada406749.

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Mossman, Jason B., and David R. Perkins. Rocket Propulsion Technology Impact on TSTO Launch System Cost. Fort Belvoir, VA: Defense Technical Information Center, May 2001. http://dx.doi.org/10.21236/ada411282.

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Saul, W., and Mark C. Grubelich. Rocket Engine Test System for Development of Novel Propulsion Technologies. Office of Scientific and Technical Information (OSTI), September 2016. http://dx.doi.org/10.2172/1562423.

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Blair, M., and D. DeGeorge. Overview of the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) Program. Fort Belvoir, VA: Defense Technical Information Center, October 2000. http://dx.doi.org/10.21236/ada411290.

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Stewart, Jesse F., and James A. Martin. Dual Fuel Solar Thermal Propulsion for LEO to GEO Transfer: Ideal Rocket Analysis. Fort Belvoir, VA: Defense Technical Information Center, July 1995. http://dx.doi.org/10.21236/ada409786.

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Geisler, R., and C. Beckman. The History of the BATES Motors at the Air Force Rocket Propulsion Laboratory. Fort Belvoir, VA: Defense Technical Information Center, July 1998. http://dx.doi.org/10.21236/ada405742.

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