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Considerations and Simulations About Pulse Detonation Engine
MATEC Web of Conferences 290, 04009 (2019) https://doi.org/10.1051/matecconf /2019290 04009 MSE 2019 Considerations and simulations about Pulse Detonation Engine Vasile Prisacariu1,*, Constantin Rotaru1, and Mihai Leonida Niculescu2 1Henri Coandă Air Force Academy, Aviation Department, Brașov, Romania 2INCAS, Aerodynamic Department, București, Romania Abstract. PDE propulsion can work from a subsonic regime to hypersonic regimes; this type of engine can have higher thermodynamic efficiency compared to other turbojet or turbofan engines due to the removal of rotating construction elements (compressors and turbines) that can reduce the mass and total cost of propulsion system. The PDE experimental researches focused on both the geometric configuration and the thermo- gas-dynamic flow aspects to prevent uncontrolled self-ignition. This article presents a series of numerical simulations on the functioning of PDE with hydrogen at supersonic regimens. Acronyms and symbols PDE pulse detonation engine AB PDE after burner pulse detonation engine DDT deflagration-to-detonation transition PDRE pulse detonation rocket engine ZDN Zel'dovich-von Neumann-Doering Mx, My Mach numbers p pressure v volume T temperature ρ density u1, u2 velocity h1, h2 enthalpy s1, s2 (S) entropy Vo,V∞ speed 1 INTRODUCTION Detonation is an effective means of burning a fuel mixture and transforming chemical energy into mechanical energy. However, this concept of running the propulsion systems involves difficulties both in rapidly achieving the fuel mixture at high speeds and in initiating and sustaining detonation in a controlled manner. PDE differs from conventional propulsion systems in two main aspects: it generates an intermittent pulse and produces high pressure increase in the combustion chamber, which also represents a major advantage of a PDE, see Figure 1. -
Call for M5 Missions
ESA UNCLASSIFIED - For Official Use M5 Call - Technical Annex Prepared by SCI-F Reference ESA-SCI-F-ESTEC-TN-2016-002 Issue 1 Revision 0 Date of Issue 25/04/2016 Status Issued Document Type Distribution ESA UNCLASSIFIED - For Official Use Table of contents: 1 Introduction .......................................................................................................................... 3 1.1 Scope of document ................................................................................................................................................................ 3 1.2 Reference documents .......................................................................................................................................................... 3 1.3 List of acronyms ..................................................................................................................................................................... 3 2 General Guidelines ................................................................................................................ 6 3 Analysis of some potential mission profiles ........................................................................... 7 3.1 Introduction ............................................................................................................................................................................. 7 3.2 Current European launchers ........................................................................................................................................... -
The Solar Cruiser Mission: Demonstrating Large Solar Sails for Deep Space Missions
The Solar Cruiser Mission: Demonstrating Large Solar Sails for Deep Space Missions Les Johnson*, Frank M. Curran**, Richard W. Dissly***, and Andrew F. Heaton* * NASA Marshall Space Flight Center ** MZBlue Aerospace NASA Image *** Ball Aerospace Solar Sails Derive Propulsion By Reflecting Photons Solar sails use photon “pressure” or force on thin, lightweight, reflective sheets to produce thrust. NASA Image 2 Solar Sail Missions Flown (as of October 2019) NanoSail-D (2010) IKAROS (2010) LightSail-1 (2015) CanX-7 (2016) InflateSail (2017) NASA JAXA The Planetary Society Canada EU/Univ. of Surrey Earth Orbit Interplanetary Earth Orbit Earth Orbit Earth Orbit Deployment Only Full Flight Deployment Only Deployment Only Deployment Only 3U CubeSat 315 kg Smallsat 3U CubeSat 3U CubeSat 3U CubeSat 10 m2 196 m2 32 m2 <10 m2 10 m2 3 Current and Planned Solar Sail Missions CU Aerospace (2018) LightSail-2 (2019) Near Earth Asteroid Solar Cruiser (2024) Univ. Illinois / NASA The Planetary Society Scout (2020) NASA NASA Earth Orbit Earth Orbit Interplanetary L-1 Full Flight Full Flight Full Flight Full Flight In Orbit; Not yet In Orbit; Successful deployed 6U CubeSat 90 Kg Spacecraft 3U CubeSat 86 m2 >1200 m2 3U CubeSat 32 m2 20 m2 4 Near Earth Asteroid Scout The Near Earth Asteroid Scout Will • Image/characterize a NEA during a slow flyby • Demonstrate a low cost asteroid reconnaissance capability Key Spacecraft & Mission Parameters • 6U cubesat (20cm X 10cm X 30 cm) • ~86 m2 solar sail propulsion system • Manifested for launch on the Space Launch System (Artemis 1 / 2020) • 1 AU maximum distance from Earth Leverages: combined experiences of MSFC and JPL Close Proximity Imaging Local scale morphology, with support from GSFC, JSC, & LaRC terrain properties, landing site survey Target Reconnaissance with medium field imaging Shape, spin, and local environment NEA Scout Full Scale EDU Sail Deployment 6 Solar Cruiser Mission Concept Mission Profile Solar Cruiser may launch as a secondary payload on the NASA IMAP mission in October, 2024. -
Propulsion Options for the Global Precipitation Measurement Core Satellite
PROPULSION OPTIONS FOR THE GLOBAL PRECIPITATION MEASUREMENT CORE SATELLITE Eric H. Cardiff*, Gary T. Davist, and David C. FoltaS NASA Goddard Space Flight Center Greenbelt, MD 2077 1 Abstract This study was conducted to evaluate several propulsion system options for the Global Precipitation Measurement (GPM) core satellite. Orbital simulations showed clear benefits for the scientific data to be obtained at a constant orbital altitude rather than with a decayheboost approach. An orbital analysis estimated the drag force on the satellite will be 1 to 12 mN during the five-year mission. Four electric propulsion systems were identified that are able to compensate for these drag forces and maintain a circular orbit. The four systems were the UK-lO/TS and the NASA 8 cm ion engines, and the ESA RMT and RITlO EVO radio-frequency ion engines. The mass, cost, and power requirements were examined for these four systems. The systems were also evaluated for the transfer time from the initial orbit of 400 x 650 km altitude orbit to a circular 400 km orbit. The transfer times were excessive, and as a consequence a “dual” system concept (with a hydrazine monopropellant system for the orbit transfer and electric propulsion for drag compensation) was examined. Clear mass benefits were obtained with the “dual” system, but cost remains an issue because of the larger power system required for the electric propulsion system. An electrodynamic tether was also evaluated in this trade study. Introduction The propulsion system is required to perform two The Global Precipitation Measurement (GPM) primary functions. The first is to transfer the mission will be launched in late 2008 to measure satellite from the launch insertion orbit to the final the amount and type of precipitation around the circular orbit. -
Einicke Diplom.Pdf
Zum Erlangen des akademischen Grades DIPLOMINGENIEUR (Dipl.-Ing.) Betreuer: Dr. Christian Bach Verantwortlicher Hochschullehrer: Prof. Dr. techn. Martin Tajmar Tag der Einreichung: 20.04.2021 Erster Gutachter: Prof. Dr. techn. Martin Tajmar Zweiter Gutachter: Dr. Christian Bach Hiermit erkläre ich, dass ich die von mir dem Institut für Luft-und Raumfahrttechnik der Fakultät Maschinenwesen eingereichte Diplomarbeit zum Thema Mischungsverhältnis- und Brennkammerdruckregelung eines Expander-Bleed Raketentriebwerks mit Reinforcement Learning (Mixture Ratio and Combustion Chamber Pressure Control of an Expander-Bleed Rocket Engine with Reinforcement Learning) selbstständig verfasst und keine anderen als die angegebenen Quellen und Hilfsmittel benutzt sowie Zitate kenntlich gemacht habe. Berlin, 20.04.2021 Karina Einicke Contents Nomenclature iv Acronyms vii 1. Introduction 1 1.1. Motivation . .1 1.2. Objectives and Approach . .2 2. Fundamentals of Liquid Rocket Engines 3 2.1. Control Loops . .5 2.1.1. Open-Loop Control . .5 2.1.2. Closed-Loop Control . .5 2.1.3. Reusable Liquid Rocket Engine Control . .6 2.2. Control Valves . .8 2.2.1. Flow Characteristics . .9 2.2.2. Valve Types . .9 2.3. Liquid Rocket Engine Control: Historical Background . 11 2.4. Summary . 13 3. LUMEN 15 3.1. LUMEN Components . 15 3.2. Operating Points . 17 3.3. EcosimPro/ESPSS Model . 18 3.3.1. LUMEN System Analysis . 21 3.3.2. LUMEN System Validation . 26 3.4. Summary . 27 4. Reinforcement Learning 28 4.1. Fundamentals of Reinforcement Learning . 28 4.2. Reinforcement Learning Algorithms . 31 4.2.1. Model-based and Model-free Reinforcement Learning . 31 4.2.2. Policy Optimization . -
In a Comprehensive New Test, the Emdrive Fails to Generate Any Thrust 7 April 2021, by Paul M
In a comprehensive new test, the EmDrive fails to generate any thrust 7 April 2021, by Paul M. Sutter conservation of momentum has been tested countless times over centuries—in fact, that principle forms the bedrock of almost every single theory of physics. So in essence, almost every time physics is tested, so is the conservation of momentum. The results of the Eagleworks experiment were not very strong. While the team claimed to measure a thrust, it wasn't statistically significant, and appeared to be a result of "cherry-picking"—the authors watching random fluctuations and waiting for the right time to report their results. But in the spirit of scientific replication, a team at the Dresden University of Technology led by Prof. Martin Tajmar rebuilt the Eagleworks experimental The EmDrive is a hypothetical rocket that setup. proponents claim can generate thrust with no exhaust. This would violate all known physics. In And they found squat. 2016, a team at NASA's Eagleworks lab claimed to measure thrust from an EmDrive device, the news Reporting their results in the Proceedings of Space of which caused quite a stir. The latest attempt to Propulsion Conference 2020, Prof. Tajmar said, replicate the shocking results has resulted in a "We found out that the cause of the 'thrust' was a simple answer: The Eagleworks measurement was thermal effect. For our tests, we used NASAs from heating of the engine mount, not any new EmDrive configuration from White et al. (which was physics. used at the Eagleworks laboratories, because it is best documented and the results were published in The EmDrive is a relatively simple device: It's an the Journal of Propulsion and Power.) empty cavity that isn't perfectly symmetrical. -
Investigation of Condensed and Early Stage Gas Phase Hypergolic Reactions Jacob Daniel Dennis Purdue University
Purdue University Purdue e-Pubs Open Access Dissertations Theses and Dissertations Fall 2014 Investigation of condensed and early stage gas phase hypergolic reactions Jacob Daniel Dennis Purdue University Follow this and additional works at: https://docs.lib.purdue.edu/open_access_dissertations Part of the Propulsion and Power Commons Recommended Citation Dennis, Jacob Daniel, "Investigation of condensed and early stage gas phase hypergolic reactions" (2014). Open Access Dissertations. 256. https://docs.lib.purdue.edu/open_access_dissertations/256 This document has been made available through Purdue e-Pubs, a service of the Purdue University Libraries. Please contact [email protected] for additional information. i INVESTIGATION OF CONDENSED AND EARLY STAGE GAS PHASE HYPERGOLIC REACTIONS A Dissertation Submitted to the Faculty of Purdue University by Jacob Daniel Dennis In Partial Fulfillment of the Requirements for the Degree of Doctor of Philosophy December 2014 Purdue University West Lafayette, Indiana ii To my parents, Jay and Susan Dennis, who have always pushed me to be the person they know I am capable of being. Also to my wife, Claresta Dennis, who not only tolerated me but suffered along with me throughout graduate school. I love you and am so proud of you! iii ACKNOWLEDGEMENTS I would like to express my sincere gratitude to my advisor, Dr. Timothée Pourpoint, for guiding me over the past four years and helping me become the researcher that I am today. In addition I would like to thank the rest of my PhD Committee for the insight and guidance. I would also like to acknowledge the help provided by my fellow graduate students who spent time with me in the lab: Travis Kubal, Yair Solomon, Robb Janesheski, Jordan Forness, Jonathan Chrzanowski, Jared Willits, and Jason Gabl. -
Rocket Propulsion Fundamentals 2
https://ntrs.nasa.gov/search.jsp?R=20140002716 2019-08-29T14:36:45+00:00Z Liquid Propulsion Systems – Evolution & Advancements Launch Vehicle Propulsion & Systems LPTC Liquid Propulsion Technical Committee Rick Ballard Liquid Engine Systems Lead SLS Liquid Engines Office NASA / MSFC All rights reserved. No part of this publication may be reproduced, distributed, or transmitted, unless for course participation and to a paid course student, in any form or by any means, or stored in a database or retrieval system, without the prior written permission of AIAA and/or course instructor. Contact the American Institute of Aeronautics and Astronautics, Professional Development Program, Suite 500, 1801 Alexander Bell Drive, Reston, VA 20191-4344 Modules 1. Rocket Propulsion Fundamentals 2. LRE Applications 3. Liquid Propellants 4. Engine Power Cycles 5. Engine Components Module 1: Rocket Propulsion TOPICS Fundamentals • Thrust • Specific Impulse • Mixture Ratio • Isp vs. MR • Density vs. Isp • Propellant Mass vs. Volume Warning: Contents deal with math, • Area Ratio physics and thermodynamics. Be afraid…be very afraid… Terms A Area a Acceleration F Force (thrust) g Gravity constant (32.2 ft/sec2) I Impulse m Mass P Pressure Subscripts t Time a Ambient T Temperature c Chamber e Exit V Velocity o Initial state r Reaction ∆ Delta / Difference s Stagnation sp Specific ε Area Ratio t Throat or Total γ Ratio of specific heats Thrust (1/3) Rocket thrust can be explained using Newton’s 2nd and 3rd laws of motion. 2nd Law: a force applied to a body is equal to the mass of the body and its acceleration in the direction of the force. -
Study and Numerical Simulation of Unconventional Engine Technology
STUDY AND NUMERICAL SIMULATION OF UNCONVENTIONAL ENGINE TECHNOLOGY by ANJALI SHEKHAR B.E Aeronautical Engineering VTU, Karnataka, 2013 A thesis submitted in partial fulfillment of the requirements for the degree of Master of Science, Aerospace Engineering, College of Engineering and Applied Science, University of Cincinnati, Ohio 2018 Thesis Committee: Chair: Ephraim Gutmark, Ph.D. Member: Shaaban Abdallah, Ph.D. Member: Mark Turner, Sc.D. An Abstract of Study and Numerical Simulation of Unconventional Engine Technology by Anjali Shekhar Submitted to the Graduate Faculty as partial fulfillment of the requirements for the Master of Science Degree in Aerospace Engineering University of Cincinnati December 2018 The aim of this thesis is to understand the working of two unconventional aircraft propul- sion systems and to setup a two-dimensional transient simulation to analyze its operational mechanism. The air traffic has nearly increased by about 40% in past three decades and calls for alternative propulsion techniques to replace or support the current traditional propulsion methodology. In the light of current demand, the thesis draws motivation from renewed inter- est in two non-conventional propulsion techniques designed in the past and had not been given due importance due to various flaws/drawbacks associated. The thesis emphasizes on the work- ing of Von Ohains thermal compression engine and pulsejet combustors. Computational Fluid Dynamics is used in current study as it offers very high flexibility and can be modified easily to incorporate the required changes. Thermal Compression engine is a design suggested by Von Ohain in 1948. The engine works on the principle of pressure rise caused inside the engine which completely depends on the temperature of working fluid and independent of rotations per minute. -
L AUNCH SYSTEMS Databk7 Collected.Book Page 18 Monday, September 14, 2009 2:53 PM Databk7 Collected.Book Page 19 Monday, September 14, 2009 2:53 PM
databk7_collected.book Page 17 Monday, September 14, 2009 2:53 PM CHAPTER TWO L AUNCH SYSTEMS databk7_collected.book Page 18 Monday, September 14, 2009 2:53 PM databk7_collected.book Page 19 Monday, September 14, 2009 2:53 PM CHAPTER TWO L AUNCH SYSTEMS Introduction Launch systems provide access to space, necessary for the majority of NASA’s activities. During the decade from 1989–1998, NASA used two types of launch systems, one consisting of several families of expendable launch vehicles (ELV) and the second consisting of the world’s only partially reusable launch system—the Space Shuttle. A significant challenge NASA faced during the decade was the development of technologies needed to design and implement a new reusable launch system that would prove less expensive than the Shuttle. Although some attempts seemed promising, none succeeded. This chapter addresses most subjects relating to access to space and space transportation. It discusses and describes ELVs, the Space Shuttle in its launch vehicle function, and NASA’s attempts to develop new launch systems. Tables relating to each launch vehicle’s characteristics are included. The other functions of the Space Shuttle—as a scientific laboratory, staging area for repair missions, and a prime element of the Space Station program—are discussed in the next chapter, Human Spaceflight. This chapter also provides a brief review of launch systems in the past decade, an overview of policy relating to launch systems, a summary of the management of NASA’s launch systems programs, and tables of funding data. The Last Decade Reviewed (1979–1988) From 1979 through 1988, NASA used families of ELVs that had seen service during the previous decade. -
MULTIPHYSICS DESIGN and SIMULATION of a TUNGSTEN-CERMET NUCLEAR THERMAL ROCKET a Thesis by BRAD APPEL Submitted to the Office O
MULTIPHYSICS DESIGN AND SIMULATION OF A TUNGSTEN-CERMET NUCLEAR THERMAL ROCKET A Thesis by BRAD APPEL Submitted to the Office of Graduate Studies of Texas A&M University in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE August 2012 Major Subject: Nuclear Engineering Multiphysics Design and Simulation of a Tungsten-Cermet Nuclear Thermal Rocket Copyright 2012 Brad Appel ii MULTIPHYSICS DESIGN AND SIMULATION OF A TUNGSTEN-CERMET NUCLEAR THERMAL ROCKET A Thesis by BRAD APPEL Submitted to the Office of Graduate Studies of Texas A&M University in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE Approved by: Chair of Committee, Karen Vierow Committee Members, Shannon Bragg-Sitton Paul Cizmas Head of Department, Yassin Hassan August 2012 Major Subject: Nuclear Engineering iii iii ABSTRACT Multiphysics Design and Simulation of a Tungsten-Cermet Nuclear Thermal Rocket. (August 2012) Brad Appel, B.S., Purdue University Chair of Advisory Committee: Dr. Karen Vierow The goal of this research is to apply modern methods of analysis to the design of a tungsten-cermet Nuclear Thermal Rocket (NTR) core. An NTR is one of the most viable propulsion options for enabling piloted deep-space exploration. Concerns over fuel safety have sparked interest in an NTR core based on tungsten-cermet fuel. This work investigates the capability of modern CFD and neutronics codes to design a cermet NTR, and makes specific recommendations for the configuration of channels in the core. First, the best CFD practices available from the commercial package Star-CCM+ are determined by comparing different modeling options with a hot-hydrogen flow experiment. -
DESIGN and DEVELOPMENT of a 30-Ghz MICROWAVE ELECTROTHERMAL THRUSTER
The Pennsylvania State University The Graduate School College of Engineering DESIGN AND DEVELOPMENT OF A 30-GHz MICROWAVE ELECTROTHERMAL THRUSTER A Thesis in Aerospace Engineering by Erica E. Capalungan ©2011 Erica E. Capalungan Submitted in Partial Fulfillment of the Requirements for the Degree of Master of Science August 2011 The thesis of Erica E. Capalungan was reviewed and approved* by the following: Michael M. Micci Professor of Aerospace Engineering Director of Graduate Studies Thesis Advisor Sven G. Bilén Associate Professor of Engineering Design, Electrical Engineering, and Aerospace Engineering George A. Lesieutre Professor of Aerospace Engineering Head of the Department of Aerospace Engineering *Signatures are on file in the Graduate School. ii ABSTRACT Research has been conducted on the microwave electrothermal thruster at The Pennsylvania State University since the 1980’s. Each subsequent thruster incorporated modifications that resulted in improvements in thruster performance compared to previous generations. Operational frequencies evaluated thus far include 2.45 GHz, 7.5 GHz, 8 GHz, and 14.5 GHz. As each thruster increased in operational frequency, plasmas have been ignited with successively lower amounts of input power. With higher frequency and lower power requirements, the physical sizes of the thruster and the power supply have been reduced. Decreased size results in a lighter propulsion system, which is ideal for space missions. This thesis concerns the design and development of a thruster operating at 30 GHz. Electromagnetic modeling was used in the design of the thruster to determine the optimal input antenna size and length. A 2.4-mm antenna size was chosen with a length that is flush with the bottom of the cavity.