Development of a 10 Kn LOX/HTPB Hybrid Rocket Engine Through Successive Development and Testing of Scaled Prototypes
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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. -
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. -
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. -
Lesson 2: YOU're LAUNCHING a ROCKET!
Adventures in Aerospace: Lesson 2 Volunteer’s Guide Key to Curriculum Formatting: ► Volunteer Directions ■ Volunteer Notes ♦ Volunteer-led Classroom Experiments Lesson 2: YOU’RE LAUNCHING A ROCKET! ► Begin the presentation by telling the class that this is “Lesson 2: You’re Launching a Rocket!” If this is your second visit, reintroduce yourself and the program. Briefly review key concepts from the first lesson, “You’re Piloting a Plane!” If this is your first visit, here is a suggested personal introduction: “Hello, my name is _____________, and I am a _________________ (position title) at Aerojet. I or another Aerojet volunteer will be visiting your class once over the next few months to speak to you about space exploration and space travel. We will learn about the basics of aerodynamics, rocket propulsion, and spaceflight to the space station, the moon, and future missions to Mars!” ► Answer any questions left over from the previous visit. MATERIALS NEEDED • AiA Multimedia Presentation (AMP) • DVD-ROM Page 1 of 11 Adventures in Aerospace: Lesson 2 Volunteer’s Guide • TV or projection screen • Handouts • Index cards ► See lesson to assess total equipment needs. LESSON OUTLINE Introduction Lesson Concepts Vocabulary Rockets vs. Airplanes Newton’s Laws • First Law • Second Law • Third Law What Type of Rocket Are You Launching? • Types of Engines Comparing and Contrasting Liquid Engines and Solid Motors Other Types of Rocket Engines Applying What We’ve Learned Experiment INTRODUCTION Rocket launches have mesmerized audiences, often entire nations, for centuries. What kind of power does it take to propel spacecraft out of the atmosphere and into the vacuum of space? This unit introduces you to rocket propulsion systems. -
Iac-15-C4.3.1 Jet Inducer for a Turbo Pump of a Liquid Rocket Engine
66th International Astronautical Congress, Jerusalem, Israel. Copyright ©2015 by the International Astronautical Federation. All rights reserved. IAC-15-C4.3.1 JET INDUCER FOR A TURBO PUMP OF A LIQUID ROCKET ENGINE Martin Böhle Technical University Kaiserslautern, Germany, [email protected] Wolfgang Kitsche German Aerospace Center (DLR), Germany, [email protected] Samuel Sudhof German Aerospace Center (DLR), Germany, [email protected] Liquid rocket propulsion systems are able to provide thrust at high specific impulse and at high thrust to weight ratios. One requirement to achieve this objective is a high combustion pressure which again requires a strong and efficient feed system. In large propulsion systems turbo pumps are used to obtain a high combustion pressure. In order to keep the mass of the pumps low the pumps are operated at very high rotational speed. Despite this a high suction performance without cavitation in the impeller is required. The state of the art designs use a so called inducer to guarantee good suction performance at high rotational speed. This publication proposes a non-rotating jet inducer which promises a higher performance and higher efficiency than the conventional inducer. The here proposed jet inducer completely replaces the conventional inducer. It works on the principle of the jet pump, a suction jet is injected before the inlet of the turbo pump and increases the total pressure of the fluid before it enters the impeller. The flow for the jet is taken from the pressure side of the pump itself and can be adjusted by a control valve depending on the operational point of the pump and on the operation phase (steady state or transient). -
Materials for Liquid Propulsion Systems
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. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. -
A Status Report Lockheed Launch Vehicle
I A STATUS REPORT LOCKHEED LAUNCH VEHICLE I for the 7th ANNUAL AIM-UTAH STATE UNIVERSITY I CONFERENCE ON SMALL SATEWTES D.E. Davis, J.W. AngelP, A.J. MaeLaren2 I Lockheed Missiles & Spaee Co.,lne Abstract: This paper discusses a new collocated product development team family of small and medium space launch with the mandate to use our engineering, vehicles being developed by Lockheed manufacturing, and flight test I Missiles & Space Company, Inc. The experience, but - and an important but development program will culminate in - to be innovative in our approaches. a demonstration launch in November For example we elected to use, as much I 1994. The paper gives a brief as possible, existing hardware; to use background and gives the program status aluminum structure rather than as of the date of this paper. composites - although Lockheed builds Supporting graphics are included. much composite structure; and to I simplify the way we do business. Background: We have streamlined our Lockheed Missiles & Space Co, paperwork system. The product I Inc. has been studying small space development team has informal reviews launch vehicles since 1987. The - informal in the sense that their are no original approach then was to use dry-runs, and hand drawn graphics are I surplus or excess ballistic missile acceptable. The technical caliber of assets, primarily the 1st and 2nd stages the presentations is as professional as if from the Poseidon C-3 missile as the it were our traditional government basis, replacing the weapon system customer instead of just ourselves. I guidance and control with state of the art Rather than tell our prospective technology, and adding a 3rd stage with subcontractors how they should design a Star 48 solid rocket motor.