Solid Rocket Propulsion for Small-Satellite Applications I C
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The SKYLON Spaceplane
The SKYLON Spaceplane Borg K.⇤ and Matula E.⇤ University of Colorado, Boulder, CO, 80309, USA This report outlines the major technical aspects of the SKYLON spaceplane as a final project for the ASEN 5053 class. The SKYLON spaceplane is designed as a single stage to orbit vehicle capable of lifting 15 mT to LEO from a 5.5 km runway and returning to land at the same location. It is powered by a unique engine design that combines an air- breathing and rocket mode into a single engine. This is achieved through the use of a novel lightweight heat exchanger that has been demonstrated on a reduced scale. The program has received funding from the UK government and ESA to build a full scale prototype of the engine as it’s next step. The project is technically feasible but will need to overcome some manufacturing issues and high start-up costs. This report is not intended for publication or commercial use. Nomenclature SSTO Single Stage To Orbit REL Reaction Engines Ltd UK United Kingdom LEO Low Earth Orbit SABRE Synergetic Air-Breathing Rocket Engine SOMA SKYLON Orbital Maneuvering Assembly HOTOL Horizontal Take-O↵and Landing NASP National Aerospace Program GT OW Gross Take-O↵Weight MECO Main Engine Cut-O↵ LACE Liquid Air Cooled Engine RCS Reaction Control System MLI Multi-Layer Insulation mT Tonne I. Introduction The SKYLON spaceplane is a single stage to orbit concept vehicle being developed by Reaction Engines Ltd in the United Kingdom. It is designed to take o↵and land on a runway delivering 15 mT of payload into LEO, in the current D-1 configuration. -
Orion Capsule Launch Abort System Analysis
Orion Capsule Launch Abort System Analysis Assignment 2 AE 4802 Spring 2016 – Digital Design and Manufacturing Georgia Institute of Technology Authors: Tyler Scogin Michel Lacerda Jordan Marshall Table of Contents 1. Introduction ......................................................................................................................................... 4 1.1 Mission Profile ............................................................................................................................. 7 1.2 Literature Review ........................................................................................................................ 8 2. Conceptual Design ............................................................................................................................. 13 2.1 Design Process ........................................................................................................................... 13 2.2 Vehicle Performance Characteristics ......................................................................................... 15 2.3 Vehicle/Sub-Component Sizing ................................................................................................. 15 3. Vehicle 3D Model in CATIA ................................................................................................................ 22 3.1 3D Modeling Roles and Responsibilities: .................................................................................. 22 3.2 Design Parameters and Relations:............................................................................................ -
Materials for Liquid Propulsion Systems
https://ntrs.nasa.gov/search.jsp?R=20160008869 2019-08-29T17:47:59+00:00Z CHAPTER 12 Materials for Liquid Propulsion Systems John A. Halchak Consultant, Los Angeles, California James L. Cannon NASA Marshall Space Flight Center, Huntsville, Alabama Corey Brown Aerojet-Rocketdyne, West Palm Beach, Florida 12.1 Introduction Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton’s third law: for every action there is an equal and opposite reaction. [1] Solid rocket motors are cheaper to manufacture and offer good values for their cost. -
Interstellar Probe on Space Launch System (Sls)
INTERSTELLAR PROBE ON SPACE LAUNCH SYSTEM (SLS) David Alan Smith SLS Spacecraft/Payload Integration & Evolution (SPIE) NASA-MSFC December 13, 2019 0497 SLS EVOLVABILITY FOUNDATION FOR A GENERATION OF DEEP SPACE EXPLORATION 322 ft. Up to 313ft. 365 ft. 325 ft. 365 ft. 355 ft. Universal Universal Launch Abort System Stage Adapter 5m Class Stage Adapter Orion 8.4m Fairing 8.4m Fairing Fairing Long (Up to 90’) (up to 63’) Short (Up to 63’) Interim Cryogenic Exploration Exploration Exploration Propulsion Stage Upper Stage Upper Stage Upper Stage Launch Vehicle Interstage Interstage Interstage Stage Adapter Core Stage Core Stage Core Stage Solid Solid Evolved Rocket Rocket Boosters Boosters Boosters RS-25 RS-25 Engines Engines SLS Block 1 SLS Block 1 Cargo SLS Block 1B Crew SLS Block 1B Cargo SLS Block 2 Crew SLS Block 2 Cargo > 26 t (57k lbs) > 26 t (57k lbs) 38–41 t (84k-90k lbs) 41-44 t (90k–97k lbs) > 45 t (99k lbs) > 45 t (99k lbs) Payload to TLI/Moon Launch in the late 2020s and early 2030s 0497 IS THIS ROCKET REAL? 0497 SLS BLOCK 1 CONFIGURATION Launch Abort System (LAS) Utah, Alabama, Florida Orion Stage Adapter, California, Alabama Orion Multi-Purpose Crew Vehicle RL10 Engine Lockheed Martin, 5 Segment Solid Rocket Aerojet Rocketdyne, Louisiana, KSC Florida Booster (2) Interim Cryogenic Northrop Grumman, Propulsion Stage (ICPS) Utah, KSC Boeing/United Launch Alliance, California, Alabama Launch Vehicle Stage Adapter Teledyne Brown Engineering, California, Alabama Core Stage & Avionics Boeing Louisiana, Alabama RS-25 Engine (4) -
Delta IV Parker Solar Probe Mission Booklet
A United Launch Alliance (ULA) Delta IV Heavy what is the source of high-energy solar particles. MISSION rocket will deliver NASA’s Parker Solar Probe to Parker Solar Probe will make 24 elliptical orbits an interplanetary trajectory to the sun. Liftoff of the sun and use seven flybys of Venus to will occur from Space Launch Complex-37 at shrink the orbit closer to the sun during the Cape Canaveral Air Force Station, Florida. NASA seven-year mission. selected ULA’s Delta IV Heavy for its unique MISSION ability to deliver the necessary energy to begin The probe will fly seven times closer to the the Parker Solar Probe’s journey to the sun. sun than any spacecraft before, a mere 3.9 million miles above the surface which is about 4 OVERVIEW The Parker percent the distance from the sun to the Earth. Solar Probe will At its closest approach, Parker Solar Probe will make repeated reach a top speed of 430,000 miles per hour journeys into the or 120 miles per second, making it the fastest sun’s corona and spacecraft in history. The incredible velocity trace the flow of is necessary so that the spacecraft does not energy to answer fall into the sun during the close approaches. fundamental Temperatures will climb to 2,500 degrees questions such Fahrenheit, but the science instruments will as why the solar remain at room temperature behind a 4.5-inch- atmosphere is thick carbon composite shield. dramatically Image courtesy of NASA hotter than the The mission was named in honor of Dr. -
National Reconnaissance Office Review and Redaction Guide
NRO Approved for Release 16 Dec 2010 —Tep-nm.T7ymqtmthitmemf- (u) National Reconnaissance Office Review and Redaction Guide For Automatic Declassification Of 25-Year-Old Information Version 1.0 2008 Edition Approved: Scott F. Large Director DECL ON: 25x1, 20590201 DRV FROM: NRO Classification Guide 6.0, 20 May 2005 NRO Approved for Release 16 Dec 2010 (U) Table of Contents (U) Preface (U) Background 1 (U) General Methodology 2 (U) File Series Exemptions 4 (U) Continued Exemption from Declassification 4 1. (U) Reveal Information that Involves the Application of Intelligence Sources and Methods (25X1) 6 1.1 (U) Document Administration 7 1.2 (U) About the National Reconnaissance Program (NRP) 10 1.2.1 (U) Fact of Satellite Reconnaissance 10 1.2.2 (U) National Reconnaissance Program Information 12 1.2.3 (U) Organizational Relationships 16 1.2.3.1. (U) SAF/SS 16 1.2.3.2. (U) SAF/SP (Program A) 18 1.2.3.3. (U) CIA (Program B) 18 1.2.3.4. (U) Navy (Program C) 19 1.2.3.5. (U) CIA/Air Force (Program D) 19 1.2.3.6. (U) Defense Recon Support Program (DRSP/DSRP) 19 1.3 (U) Satellite Imagery (IMINT) Systems 21 1.3.1 (U) Imagery System Information 21 1.3.2 (U) Non-Operational IMINT Systems 25 1.3.3 (U) Current and Future IMINT Operational Systems 32 1.3.4 (U) Meteorological Forecasting 33 1.3.5 (U) IMINT System Ground Operations 34 1.4 (U) Signals Intelligence (SIGINT) Systems 36 1.4.1 (U) Signals Intelligence System Information 36 1.4.2 (U) Non-Operational SIGINT Systems 38 1.4.3 (U) Current and Future SIGINT Operational Systems 40 1.4.4 (U) SIGINT -
Development of a 10 Kn LOX/HTPB Hybrid Rocket Engine Through Successive Development and Testing of Scaled Prototypes
DOI: 10.13009/EUCASS2017-334 7TH EUROPEAN CONFERENCE FOR AERONAUTICS AND AEROSPACE SCIENCES (EUCASS) Development of a 10 kN LOX/HTPB Hybrid Rocket Engine through Successive Development and Testing of Scaled Prototypes Maximilian Bambauer and Markus Brandl WARR (TUM) c/o TUM Lehrstuhl für Raumfahrttechnik Boltzmannstraße 15 D-85748 Garching bei München [email protected] [email protected] · Abstract The development of a hybrid flight engine for a sounding rocket requires intensive preparation work and pretests. To obtain better understanding of the difficulties in design and operation that occur with LOX/HTPB rocket engines, four rocket engines with increasing size and thrust level are successively developed and manufactured. Listed chronologically, those engines are the demonstrator engine, the sub- scale engine with a cylindrical grain geometry and later a star shaped geometry, the full-scale test engine and the full-scale flight version. At first a small-scale technology demonstrator was developed. It pro- duced a mean thrust of 160N for 5s and had the purpose to validate the design calculations and to gain first experiences with the use of cryogenic propellants, especially regarding cooldown procedures. Based on this test results the so-called sub-scale engine was developed and tested. The engine can be operated with two grain configurations, using either a cylindrical grain or a star shaped grain. The star shaped grain geometries are realized, using an additive manufactured core, outside of which the HTPB is casted. In the low thrust configuration, the engine is operated using a cylindrical single port HTPB grain and it produces a thrust of about 540 N for a duration of 10 s. -
The Delta Launch Vehicle- Past, Present, and Future
The Space Congress® Proceedings 1981 (18th) The Year of the Shuttle Apr 1st, 8:00 AM The Delta Launch Vehicle- Past, Present, and Future J. K. Ganoung Manager Spacecraft Integration, McDonnell Douglas Astronautics Co. H. Eaton Delta Launch Program, McDonnell Douglas Astronautics Co. Follow this and additional works at: https://commons.erau.edu/space-congress-proceedings Scholarly Commons Citation Ganoung, J. K. and Eaton, H., "The Delta Launch Vehicle- Past, Present, and Future" (1981). The Space Congress® Proceedings. 7. https://commons.erau.edu/space-congress-proceedings/proceedings-1981-18th/session-6/7 This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. THE DELTA LAUNCH VEHICLE - PAST, PRESENT AND FUTURE J. K. Ganoung, Manager H. Eaton, Jr., Director Spacecraft Integration Delta Launch Program McDonnell Douglas Astronautics Co. McDonnell Douglas Astronautics Co. INTRODUCTION an "interim space launch vehicle." The THOR was to be modified for use as the first stage, the The Delta launch vehicle is a medium class Vanguard second stage propulsion system, was used expendable booster managed by the NASA Goddard as the Delta second stage and the Vanguard solid Space Flight Center and used by the U.S. rocket motor became Delta's third stage. Government, private industry and foreign coun Following the eighteen month development program tries to launch scientific, meteorological, and failure to launch its first payload into or applications and communications satellites. -
A Tool for Preliminary Design of Rockets Aerospace Engineering
A Tool for Preliminary Design of Rockets Diogo Marques Gaspar Thesis to obtain the Master of Science Degree in Aerospace Engineering Supervisor : Professor Paulo Jorge Soares Gil Examination Committee Chairperson: Professor Fernando José Parracho Lau Supervisor: Professor Paulo Jorge Soares Gil Members of the Committee: Professor João Manuel Gonçalves de Sousa Oliveira July 2014 ii Dedicated to my Mother iii iv Acknowledgments To my supervisor Professor Paulo Gil for the opportunity to work on this interesting subject and for all his support and patience. To my family, in particular to my parents and brothers for all the support and affection since ever. To my friends: from IST for all the companionship in all this years and from Coimbra for the fellowship since I remember. To my teammates for all the victories and good moments. v vi Resumo A unica´ forma que a humanidade ate´ agora conseguiu encontrar para explorar o espac¸o e´ atraves´ do uso de rockets, vulgarmente conhecidos como foguetoes,˜ responsaveis´ por transportar cargas da Terra para o Espac¸o. O principal objectivo no design de rockets e´ diminuir o peso na descolagem e maximizar o payload ratio i.e. aumentar a capacidade de carga util´ ao seu alcance. A latitude e o local de lanc¸amento, a orbita´ desejada, as caracter´ısticas de propulsao˜ e estruturais sao˜ constrangimentos ao projecto do foguetao.˜ As trajectorias´ dos foguetoes˜ estao˜ permanentemente a ser optimizadas, devido a necessidade de aumento da carga util´ transportada e reduc¸ao˜ do combust´ıvel consumido. E´ um processo utilizado nas fases iniciais do design de uma missao,˜ que afecta partes cruciais do planeamento, desde a concepc¸ao˜ do ve´ıculo ate´ aos seus objectivos globais. -
ITEM 2 Complete Subsystems Complete
ITEM 2 Complete Subsystems Complete CATEGORY I ~ ITEM 2 CATEGORY Subsystems Complete subsystems usable in the systems in Item 1, as follows, as well Produced by as the specially designed “production facilities” and “production equip- companies in ment” therefor: (a) Individual rocket stages; • Brazil • China Nature and Purpose: A rocket stage generally consists of structure, engine/ • Egypt motor, propellant, and some elements of a control system. Rocket engines or • France motors produce propulsive thrust to make the rocket fly by using either solid • Germany propellants, which burn to exhaustion once ignited, or liquid propellants • India burned in a combustion chamber fed by pressure tanks or pumps. • Iran • Iraq Method of Operation: A launch signal either fires an igniter in the solid • Israel propellant inside the lowermost (first stage) rocket motor, or tank pressure • Italy or a pump forces liquid propellants into the combustion chamber of a • Japan rocket engine where they react. Expanding, high-temperature gases escape • Libya at high speeds through a nozzle at the rear of the rocket stage. The mo- • North Korea mentum of the exhausting gases provides the thrust for the missile. Multi- • Pakistan stage rocket systems discard the lower stages as they burn up their propel- • Russia lant and progressively lose weight, thereby achieving greater range than • South Korea comparably sized, single-stage rocket systems. • Syria • Ukraine Typical Missile-Related Uses: Rocket stages are necessary components of any • United Kingdom rocket system. Rocket stages also are used in missile and missile-component • United States testing applications. Photo Credit: Chemical Propulsion Information Agency Other Uses: N/A Appearance (as manufactured): Solid propellant rocket motor stages are cylinders usually ranging from 4 to 10 m in length and 0.5 to 4 m in diame- ter, and capped at each end with hemispherical domes as shown in Figure 2-1. -
DESIGN, of MULTI-PROPELLANT STAR GRAINS for SOLID PROPELLANT ROCKETS S. KRISHNAN & TARIT K. BOSB When Simplicity, Reliabilit
'c DESIGN, OF MULTI-PROPELLANT STAR GRAINS FOR SOLID PROPELLANT ROCKETS / S. KRISHNAN& TARIT K. BOSB Department of Aeronautical Engineering Indian Institute of Technology, Madras-600036 (Received 29 November 1976; revised 24 July 1979) A new approach to solve the geometry-problem of solid propellant star grains is presented. The basis of the approach is to take the web-thickness (a ballistic as well as a geomstrical propsrty) ar th: chpracteristic length. The nondimensional characteristicparameters representing diamster, length, slen3erne;s-ratio, and ignitor accomnoda- tion of the grain are all identified. Many particular cases of star configuratiols (from the c~nfigurations of single propellant to those of four different propellants) can be analysed through the identified characteristic para- meters. A better way of representing the s~ngle-propellant-star-performancein a design graph is explained. Two types of dual propella+ grains are analysed in detail. The tirst type is characterised by its two distinct stages of burning (initially by slngle propellant burning and then by dual propellant burning); the second type has the dual propellant burning throughout. Sultab~btyof the identified charactenstic paraaeters to an optimisation study is demonstrated through examples. When simplicity, reliability, development time and cost are the most important factors, the star con- figuration is one of the commonly opted geometries. Typical applications of star grains include tactical , and sounding rockets, and satellite launch vehicles1-*. Eva in large high performance segmented grains, sometimes star configuration is uses for some of the segments5. Although from mechanical behaviour point of view the wagon-wheel grains are sometimes superior to star grains, they usually do not give a web as thick as star grains. -
Extensions to the Time Lag Models for Practical Application to Rocket
The Pennsylvania State University The Graduate School College of Engineering EXTENSIONS TO THE TIME LAG MODELS FOR PRACTICAL APPLICATION TO ROCKET ENGINE STABILITY DESIGN A Dissertation in Mechanical Engineering by Matthew J. Casiano © 2010 Matthew J. Casiano Submitted in Partial Fulfillment of the Requirements for the Degree of Doctor of Philosophy August 2010 The dissertation of Matthew J. Casiano was reviewed and approved* by the following: Domenic A. Santavicca Professor of Mechanical Engineering Co-chair of Committee Vigor Yang Adjunct Professor of Mechanical Engineering Dissertation Advisor Co-chair of Committee Richard A. Yetter Professor of Mechanical Engineering André L. Boehman Professor of Fuel Science and Materials Science and Engineering Tomas E. Nesman Aerospace Engineer at NASA Marshall Space Flight Center Special Member Karen A. Thole Professor of Aerospace Engineering Head of the Department of Mechanical and Nuclear Engineering *Signatures are on file in the Graduate School iii ABSTRACT The combustion instability problem in liquid-propellant rocket engines (LREs) has remained a tremendous challenge since their discovery in the 1930s. Improvements are usually made in solving the combustion instability problem primarily using computational fluid dynamics (CFD) and also by testing demonstrator engines. Another approach is to use analytical models. Analytical models can be used such that design, redesign, or improvement of an engine system is feasible in a relatively short period of time. Improvements to the analytical models can greatly aid in design efforts. A thorough literature review is first conducted on liquid-propellant rocket engine (LRE) throttling. Throttling is usually studied in terms of vehicle descent or ballistic missile control however there are many other cases where throttling is important.