DOCSLIB.ORG

A Tool for Preliminary Design of Rockets Aerospace Engineering

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text

Load more

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. O principal objectivo deste trabalho e´ encontrar o design mais adequado de um foguetao,˜ para uma orbita´ e carga espec´ıficas, considerando a melhor trajectoria´ que pode ser utilizada. Foi desenvolvido um programa que permite o estudo de diferentes configurac¸oes˜ e a variac¸ao˜ dos parametrosˆ de design, calculando para cada configurac¸ao˜ e conjunto de parametrosˆ as diferentes massas e dimensoes˜ do foguetao,˜ para depois iterativamente com a trajectoria´ conseguir maximizar o payload ratio, minimizando o peso na descolagem. Para validar o codigo´ foram feitas comparac¸oes˜ com os resultados obtidos para os lanc¸adores Vega e Proton K/DM3. Por ultimo,´ e´ feita uma optimizac¸ao˜ do lanc¸ador Ariane 5 atraves´ da variac¸ao˜ da sua configurac¸ao˜ e de parametrosˆ de design. Palavras-chave: Multi-Estagios´ Rockets, Gravity Turn, Optimizac¸ao˜ de Trajectoria,´ Optimizac¸ao˜ de Rockets, Maximizar o Payload Ratio. vii viii Abstract The only way mankind can explore space is with the use of space launch vehicles, commonly known as rockets, which carry payloads from Earth into Space. The main goal in the rocket design is to reduce the gross lift-off weight (GLOW) and increase the payload ratio, giving them the capacity of having the maximum amount of payload or its range. Launch latitude, launch site, final kick-off location for payload, propulsion characteristics and, or, its own struc- tural characteristics are all limitations in the launch vehicle’s design. The trajectories of launch vehicles are permanently being optimized, due to the demand of increased payload and reduced propellant requirements. Optimization of a trajectory is a process used in early stages of the mission design, it affects crucial parts of mission planning, ranging from vehicle design to even overall mission objectives. The main objective of this study is to find the best design of a rocket for a specific orbit and payload mass, considering the best trajectory that can be used. A tool was developed allowing the study of different configurations and the variation of design parameters, calculating for each configuration and parameters the different masses and dimensions that characterize a rocket, and then iteratively with the trajectory will achieve the minimum GLOW and obtain the maximum payload ratio. In order to validate the code comparisons with Vega and Proton K/DM3 launchers were made. At last an optimization of the Ariane 5 launcher is made by varying its configuration and some design parameters. Keywords: Multistage Rockets, Gravity Turn, Trajectory optimization, Rocket optimization, Max- imize Payload Ratio. ix x Contents Acknowledgments...........................................v Resumo................................................. vii Abstract................................................. ix List of Tables.............................................. xv List of Figures............................................. xviii Nomenclature.............................................. xx Glossary................................................ xxi 1 Introduction 1 1.1 Goal of this work.........................................1 1.2 History of Rockets........................................1 1.3 Challenges in the design of a Rocket..............................3 1.4 Work overview..........................................5 2 Rockets 7 2.1 Introduction to Rockets......................................7 2.2 Delta-V calculations.......................................7 2.3 Configuration...........................................8 2.3.1 Multi Stage........................................8 2.3.2 Boosters......................................... 10 2.4 Structure............................................. 11 2.4.1 Lower Stages....................................... 11 2.4.2 Upper Stages....................................... 11 2.4.3 Fairing........................................... 12 2.4.4 Other components.................................... 13 2.5 Propulsion............................................. 13 2.5.1 Solid Propellants..................................... 14 2.5.2 Liquid Propellants.................................... 15 2.5.3 Hybrid Propellants.................................... 16 3 Ascent Rocket Trajectories 19 3.1 Trajectories............................................ 19 xi 3.1.1 Ascent Phase....................................... 19 3.1.2 Gravity Turn........................................ 20 3.1.3 Free-flight phase..................................... 21 3.1.4 Coast Phases....................................... 24 3.1.5 Drag Model........................................ 25 3.1.6 Atmospheric Model................................... 26 3.1.7 Gravitational Model................................... 27 4 Rocket Knowledge Database 29 4.1 DB construction development.................................. 29 4.1.1 Small........................................... 29 4.1.2 Medium.......................................... 29 4.1.3 Heavy........................................... 30 4.2 Information gathered....................................... 30 4.3 Propellant Selection....................................... 33 4.4 Mass Estimation Relationships................................. 34 5 Rocket Design Tool 37 5.1 Program Overview........................................ 37 5.2 Parametrization.......................................... 38 5.3 Algorithm............................................. 38 5.3.1 Mass Model........................................ 40 5.3.2 Trajectory......................................... 42 6 Results 45 6.1 Validation............................................. 45 6.1.1 Trajectory......................................... 45 6.1.2 Mass model validation.................................. 47 6.2 Preliminary Design and Optimization.............................. 53 7 Conclusions 59 7.1 Future Work............................................ 61 Bibliography 67 A Nose Cone Geometries 69 A.1 Nose Cone Geometries..................................... 69 A.1.1 Ogive........................................... 69 A.1.2 Power........................................... 70 A.1.3 Ellipse........................................... 71 A.1.4 Haack........................................... 71 xii B Propellant properties 73 xiii xiv List of Tables 4.1 Characteristics gathered for each Stage of each Launcher.................. 30 4.2 Structural factor for each Stage................................. 31 4.3 Thrust in Vacuum for each Stage................................ 31 4.4 Length/Diameter ratio...................................... 31 4.5 Characteristics of Payload Fairings............................... 32 4.6 Ratio between Mass Fairing and GLOW............................ 32 4.7 Propellant Specifications [1]................................... 33 5.1 Configuration and parameters introduced in the tool...................... 38 5.2 Parameters of the Exterior Structure [2][3]........................... 41 6.1 Vega Launcher - Main characteristics [4] ............................ 46 6.2 Proton K/DM3 - Main characteristics [5]. ............................ 46 6.3 Configuration and parameters introduced in the tool..................... 47 6.4 Mass Model Comparison - Vega ................................ 48 6.5 Volume Comparison - Vega ................................... 48 6.6 Mass Model Comparison - PROTON K ............................ 50 6.7 Volume Comparison - Proton K ................................. 51 6.8 Ariane 5 Launcher - Main characteristics. ........................... 53 6.9 Parameters varied in the tool................................... 54 6.10 Optimal Configuration Achieved................................. 55 6.11 Mass Model Comparison - Ariane 5 .............................. 55 6.12 Volume Comparison - Ariane 5 ................................
Recommended publications
  • Space Administration

    Space Administration

    https://ntrs.nasa.gov/search.jsp?R=19700024651 2020-03-23T18:20:34+00:00Z TO THE CONGRESSOF THE UNITEDSTATES : Transmitted herewith is the Twenty-first Semiannual Repol* of the National Aeronautics and Space Administration. Twen~-first SEMIANNUAL REPORT TO CONGRESS JANUARY 1 - JUNE 30, 1969 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. 20546 Editors: G. B. DeGennaro, H. H. Milton, W. E. Boardman, Office of Public Affairs; Art work: A. Jordan, T. L. Lindsey, Office of Organiza- tion and Management. For sale by the Superintendent of Documents, U.S. Government Printing Office Washington, D.C. 20402-Price $1.25 THE PRESIDENT May 27,1970 The White House I submit this Twenty-First Semiannual Report of the National Aeronautics and Space Aldministration to you for transmitttal to Congress in accordance with section 206(a) of the National Aero- nautics and Space Act of 1958. It reports on aotivities which took place betiween January 1 and June 30, 1969. During this time, the Nation's space program moved forward on schedule. ApolIo 9 and 10 demonstrated the ability of ;the man- ned Lunar Module to operate in earth and lunar orbit and its 'eadi- ness to attempt the lunar landing. Unmanned observatory and ex- plorer class satellites carried on scientific studies of the regions surrounding the Earth, the Moon, and the Sun; a Biosatellite oarwing complex biological science experiment was orbited; and sophisticated weather satellites and advanced commercial com- munications spacecraft became operational. Advanced research projects expanded knowledge of space flighk and spacecraft engi- neering as well as of aeronautics.
  • Progress Report on Apollo Program

    Progress Report on Apollo Program

    PROGRESS REPORT ON APOLLO PROGRAM Michael Collins, LCol. USAF (M) Astronaut NASA-MANNED SPACECRAFT CENTER It is a great pleasure to be here today and to greet you hardy suMvors of the pool party. I will do my best to avoid loud noises and bright colors during my status report. Since the last SETP Symposium, the Apollo Program has been quite busy in a number of different areas. (Figure 1) My problem is to sift through this information and to talk only about those things of most interest to you. First, to review briefly our hardware, we are talking about two different spacecraft and two different boosters. (Figure 2) The Command Module is that part of the stack COLLINS which makes the complete round trip to the moon. Attached to it is the Service Module, containing expendables and a 20,000 pound thrust engine for maneuverability. The Lunar Module will be carried on later flights and is the landing vehicle and active rendezvous partner. The uprated Saturn I can put the Command and Service Modules into earth orbit; the Saturn V is required when the Lunar Module is added. Since the last symposium, we have flown the Command and Service Modules twice and the Lunar Module once, all unmanned. Apollo 4, the first Saturn V flight, was launched in November 1967. (Figure 3) The Saturn V did a beautiful, i.e. nominal, job of putting the spacecraft into earth parking orbit. After a coast period, the third stage (S-IVB by McDonnell Douglas) was ignited a second time, achieving a highly elliptical orbit.
  • The SKYLON Spaceplane

    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.
  • Rocket Propulsion Fundamentals 2

    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.
  • Coplanar Air Launch with Gravity-Turn Launch Trajectories

    Coplanar Air Launch with Gravity-Turn Launch Trajectories

    View metadata, citation and similar papers at core.ac.uk brought to you by CORE provided by AFTI Scholar (Air Force Institute of Technology) Air Force Institute of Technology AFIT Scholar Theses and Dissertations Student Graduate Works 3-23-2004 Coplanar Air Launch with Gravity-Turn Launch Trajectories David W. Callaway Follow this and additional works at: https://scholar.afit.edu/etd Part of the Aerospace Engineering Commons Recommended Citation Callaway, David W., "Coplanar Air Launch with Gravity-Turn Launch Trajectories" (2004). Theses and Dissertations. 3922. https://scholar.afit.edu/etd/3922 This Thesis is brought to you for free and open access by the Student Graduate Works at AFIT Scholar. It has been accepted for inclusion in Theses and Dissertations by an authorized administrator of AFIT Scholar. For more information, please contact [email protected]. COPLANAR AIR LAUNCH WITH GRAVITY-TURN LAUNCH TRAJECTORIES THESIS David W. Callaway, 1st Lieutenant, USAF AFIT/GAE/ENY/04-M04 DEPARTMENT OF THE AIR FORCE AIR UNIVERSITY AIR FORCE INSTITUTE OF TECHNOLOGY Wright-Patterson Air Force Base, Ohio APPROVED FOR PUBLIC RELEASE; DISTRIBUTION UNLIMITED. The views expressed in this thesis are those of the author and do not reflect the official policy or position of the United States Air Force, Department of Defense, or the United States Government. AFIT/GAE/ENY/04-M04 COPLANAR AIR LAUNCH WITH GRAVITY-TURN LAUNCH TRAJECTORIES THESIS Presented to the Faculty Department of Aeronautics and Astronautics Graduate School of Engineering and Management Air Force Institute of Technology Air University Air Education and Training Command In Partial Fulfillment of the Requirements for the Degree of Master of Science in Aeronautical Engineeering David W.
  • Mscthesis Joseangelgutier ... Humada.Pdf

    Mscthesis Joseangelgutier ... Humada.Pdf

    Targeting a Mars science orbit from Earth using Dual Chemical-Electric Propulsion and Ballistic Capture Jose Angel Gutierrez Ahumada Delft University of Technology TARGETING A MARS SCIENCE ORBIT FROM EARTH USING DUAL CHEMICAL-ELECTRIC PROPULSION AND BALLISTIC CAPTURE by Jose Angel Gutierrez Ahumada MSc Thesis in partial fulfillment of the requirements for the degree of Master of Science in Aerospace Engineering at the Delft University of Technology, to be defended publicly on Wednesday May 8, 2019. Supervisor: Dr. Francesco Topputo, TU Delft/Politecnico di Milano Dr. Ryan Russell, The University of Texas at Austin Thesis committee: Dr. Francesco Topputo, TU Delft/Politecnico di Milano Prof. dr. ir. Pieter N.A.M. Visser, TU Delft ir. Ron Noomen, TU Delft Dr. Angelo Cervone, TU Delft This thesis is confidential and cannot be made public until December 31, 2020. An electronic version of this thesis is available at http://repository.tudelft.nl/. Cover picture adapted from https://steemitimages.com/p/2gs...QbQvi EXECUTIVE SUMMARY Ballistic capture is a relatively novel concept in interplanetary mission design with the potential to make Mars and other targets in the Solar System more accessible. A complete end-to-end interplanetary mission from an Earth-bound orbit to a stable science orbit around Mars (in this case, an areostationary orbit) has been conducted using this concept. Sets of initial conditions leading to ballistic capture are generated for different epochs. The influence of the dynamical model on the capture is also explored briefly. Specific capture trajectories are then selected based on a study of their stabilization into an areostationary orbit.
  • Atlas V Cutaway Poster

    Atlas V Cutaway Poster

    ATLAS V Since 2002, Atlas V rockets have delivered vital national security, science and exploration, and commercial missions for customers across the globe including the U.S. Air Force, the National Reconnaissance Oice and NASA. 225 ft The spacecraft is encapsulated in either a 5-m (17.8-ft) or a 4-m (13.8-ft) diameter payload fairing (PLF). The 4-m-diameter PLF is a bisector (two-piece shell) fairing consisting of aluminum skin/stringer construction with vertical split-line longerons. The Atlas V 400 series oers three payload fairing options: the large (LPF, shown at left), the extended (EPF) and the extra extended (XPF). The 5-m PLF is a sandwich composite structure made with a vented aluminum-honeycomb core and graphite-epoxy face sheets. The bisector (two-piece shell) PLF encapsulates both the Centaur upper stage and the spacecraft, which separates using a debris-free pyrotechnic actuating 200 ft system. Payload clearance and vehicle structural stability are enhanced by the all-aluminum forward load reactor (FLR), which centers the PLF around the Centaur upper stage and shares payload shear loading. The Atlas V 500 series oers 1 three payload fairing options: the short (shown at left), medium 18 and long. 1 1 The Centaur upper stage is 3.1 m (10 ft) in diameter and 12.7 m (41.6 ft) long. Its propellant tanks are constructed of pressure-stabilized, corrosion-resistant stainless steel. Centaur is a liquid hydrogen/liquid oxygen-fueled vehicle. It uses a single RL10 engine producing 99.2 kN (22,300 lbf) of thrust.
  • Solid Propellant Rocket Engines - V.M

    Solid Propellant Rocket Engines - V.M

    THERMAL TO MECHANICAL ENERGY CONVERSION: ENGINES AND REQUIREMENTS – Vol. II - Solid Propellant Rocket Engines - V.M. Polyaev and V.A. Burkaltsev SOLID PROPELLANT ROCKET ENGINES V.M. Polyaev and V.A. Burkaltsev Department of Rocket engines, Bauman Moscow State Technical University, Russia. Keywords: Combustion chamber, solid propellant load, pressure, temperature, nozzle, thrust, control, ignition, cartridge, aspect ratio, regime. Contents 1. Introduction 2. Historical information 3. SPRE scheme and main units 4. SPRE operation 5. Parameter optimization, the approach and results 6. Transient regime 7. Service 8. Development prospects 9. Conclusions Acknowledgments Glossary Bibliography Biographical Sketches 1. Introduction Solid propellant rocket engines (SPRE) are called the direct reaction engines, in which chemical energy of the solid propellant being placed in the combustion chamber is transformed at first to thermal energy, and then to kinetic energy of the combustion products thrown away with high velocity in the environment. The momentum of combustion products discharging through the nozzle is equal to the impulse of reactive force being created by the engine. 2. Historical information First of rocketsUNESCO known to us were rockets – with EOLSS primitive powder rocket engines used in China near 5000 years ago for pleasure and military aims (so-called "fiery arrows", Figure 1). SAMPLE CHAPTERS The first rocket propellant was black smoky powder (potassium saltpetre with charcoal mixture). In Russia, powder rockets appeared in the beginning of XVII century. In 1680, Tsar Peter I founded "rocket institution" in Moscow for firework rockets making. In 1717 lighting signal rockets existed for 200 years without changes. In the beginning of XIX century, Englishman Kongrev improved the "fiery arrows" having been borrowed from Hindus.
  • Modeling Rocket Flight in the Low-Friction Approximation

    Modeling Rocket Flight in the Low-Friction Approximation

    Undergraduate Journal of Mathematical Modeling: One + Two Volume 6 | 2014 Fall Article 5 2014 Modeling Rocket Flight in the Low-Friction Approximation Logan White University of South Florida Advisors: Manoug Manougian, Mathematics and Statistics Razvan Teodorescu, Physics Problem Suggested By: Razvan Teodorescu Follow this and additional works at: https://scholarcommons.usf.edu/ujmm Part of the Mathematics Commons UJMM is an open access journal, free to authors and readers, and relies on your support: Donate Now Recommended Citation White, Logan (2014) "Modeling Rocket Flight in the Low-Friction Approximation," Undergraduate Journal of Mathematical Modeling: One + Two: Vol. 6: Iss. 1, Article 5. DOI: http://dx.doi.org/10.5038/2326-3652.6.1.4861 Available at: https://scholarcommons.usf.edu/ujmm/vol6/iss1/5 Modeling Rocket Flight in the Low-Friction Approximation Abstract In a realistic model for rocket dynamics, in the presence of atmospheric drag and altitude-dependent gravity, the exact kinematic equation cannot be integrated in closed form; even when neglecting friction, the exact solution is a combination of elliptic functions of Jacobi type, which are not easy to use in a computational sense. This project provides a precise analysis of the various terms in the full equation (such as gravity, drag, and exhaust momentum), and the numerical ranges for which various approximations are accurate to within 1%. The analysis leads to optimal approximations expressed through elementary functions, which can be implemented for efficient flight
  • National Reconnaissance Office Review and Redaction Guide

    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
  • Humanity and Space

    Humanity and Space

    10/17/2012!! !!!!!! Project Number: MH-1207 Humanity and Space An Interactive Qualifying Project Submitted to WORCESTER POLYTECHNIC INSTITUTE In partial fulfillment for the Degree of Bachelor of Science by: Matthew Beck Jillian Chalke Matthew Chase Julia Rugo Professor Mayer H. Humi, Project Advisor Abstract Our IQP investigates the possible functionality of another celestial body as an alternate home for mankind. This project explores the necessary technological advances for moving forward into the future of space travel and human development on the Moon and Mars. Mars is the optimal candidate for future human colonization and a stepping stone towards humanity’s expansion into outer space. Our group concluded space travel and interplanetary exploration is possible, however international political cooperation and stability is necessary for such accomplishments. 2 Executive Summary This report provides insight into extraterrestrial exploration and colonization with regards to technology and human biology. Multiple locations have been taken into consideration for potential development, with such qualifying specifications as resources, atmospheric conditions, hazards, and the environment. Methods of analysis include essential research through online media and library resources, an interview with NASA about the upcoming Curiosity mission to Mars, and the assessment of data through mathematical equations. Our findings concerning the human aspect of space exploration state that humanity is not yet ready politically and will not be able to biologically withstand the hazards of long-term space travel. Additionally, in the field of robotics, we have the necessary hardware to implement adequate operational systems yet humanity lacks the software to implement rudimentary Artificial Intelligence. Findings regarding the physics behind rocketry and space navigation have revealed that the science of spacecraft is well-established.
  • Launch Vehicle Optimization

    Launch Vehicle Optimization

    Launch vehicle optimization 1.Orbital Mechanics for engineering students Chapter – 11: Rocket vehicle dynamics 2. Space Flight Dynamics By William E Wiesel Chapter 7 – Rocket Performance Restricted staging in field-free space No gravity and no aerodynamics Let gross mass of a launch vehicle m0 = empty mass mE + propellant mass mp + payload mass mPL Empty mass mE = mass of structure + mass of fuel tank and related system + mass of control system Let us divide the above by m0 We can write as structural fraction , πE = mE / m0 Propellant fraction, πp =mp / m0 payload fraction, πPL = mPL / m0 Alternately we can define Payload ratio Structural ratio Mass ratio Assuming all the propellant Is consumed λ, ε and n are not independent From we can write as From we can write as Substituting and in We get Given any two of the ratios λ, ε and n, we can obtain the third Velocity at burn out is For a given empty mass, the greatest possible Δv occurs when the payload is zero. To maximize the amount of payload while keeping the structural weight to a minimum. Mass of load-bearing structure, rocket motors, pumps, piping, etc., cannot be made arbitrarily small. Current materials technology places a lower limit on ε of about 0.1. Performance of multistage rocket Restricted staging - all stages are similar Each stage has the same specific impulse Isp same structural ratio ε same payload ratio λ. Hence mass ratios n are identical Final burnout speed vbo for a given payload mass mPL Overall payload fraction m0 is the total mass of the tandem-stacked vehicle.