DOCSLIB.ORG

A Status Report Lockheed Launch Vehicle

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
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. Although their hardware, we have either accepted I the approach was notionally attractive their specification or we have discussed and the calculated reliability was over our requirements and allowed them to 90%, no users were found. suggest the solutions and write the Motorola's announcement of the specification. We have held an oral I Iridium™ program in 1990 competition where the proposal was the dramatically altered our approach. viewgraphs and a one page letter on Motorola's imperatives were very clear: price. This has not been without I lowest possible cost, on demand launCh, some internal trauma and there have and outstanding reliability. When we been replacements of personnel who worked through the alternatives, we could not cope with the new world order. I concluded that reliability had to be as On January 8, 1993 Lockheed close to 100% as possible. A 90% to Corporation approved a company funded 95% reliability was unacceptable; we development program and demon­ calculated that for program such as stration launch in November 1994. I Iridium™ we had to self-insure, because We have submitted the required range the costs of insurance at 15 to 18 documentation to both the Eastern Range percent per launch would be (Cape Canaveral) and the Western Range I unacceptable from a price strategy point (Vandenberg Air Force Base). The Air of view. Thus we quickly discarded the Force has approved our use of Space ballistic missile assets approach, and Launch Complex (SLC-6) at Vandenberg I reexamined approaches based on AFB. We are working with the Eastern Thiokol's Castor 120™ which was then Range and the Florida Space Port I in development. We formed a Authority at Cape Canaveral. We have I 1. P.E., Member ASME 2. Senior Member AIAA I 2 started discussions with the Office of at the launch site to final countdown and I Commercial Space Transportation launch is a short as two weeks. (OCST) of the Department of The first of these vehicles is Transportation (DoT) with respect to scheduled to fly from Vandenberg in I the relevant licenses required by the November 1994. General operations Commercial Space Launch Act (CSLA). will begin in 1995. We completed our Preliminary Design Review (PDR) process in July, Solid Rocket Motors: I and have started fabrication of the first The LL V1 first stage is Thiokol hardware designated for testing. We Corporation's Castor 120™ Solid Rocket are on a clear track to meet the Motor. It is a 120,000 Ibm class motor I November 1994 launch date mandated that employs a graphite epoxy composite by our CEO, Dan Tellep. case, Class 1.3 HTPB propellant, pyrogen igniter, and a vectorable nozzle Launch Vehicle Description: driven by a cold gas blow down thrust I This section describes the design vector control system. This motor is and system performance of this new used as both the first stage and the family of launch vehicles designated the second stage on the LLV2 and LLV3. As a I Lockheed Launch Vehicles. second stage motor, a larger expansion The smallest vehicle (LLV1) will ratio nozzle is used in order to improve place about 2200 pounds in to a low performance at higher altitudes. The I Earth orbit (LEO) of 100 nautical miles propellant grain can be tailored to at 28°. The next increment, (LLV2) change the burn characteristics and will place 4000 pounds into LEO, and thrust history. This provides the the LLV3 will place up to 8000 pounds ability to taper the thrust of the second I into LEO. Launch operations will be stage near burn out, which lowers the conducted from Vandenberg Air Force vehicle acceleration and provide~ a Base in California for Polar and Sun smoother ride for the payload. I Synchronous orbits and from Cape The upper stage of the LLV family Canaveral. A mobile launch system is the Orbus 21 D solid rocket motor checkout van will be provided which manufactured by the Chemical Systems will allow launch operations from any Division of United Technologies. Inc. I suitable range. The Orbus 21 is a proven motor which Launch Vehicle assembly and has been used in the Inertial Upper Stage checkout at the launch site is based on a on 17 flights. For use on the LLV it will I simple stack and shoot approach allowed use a carbon phenolic nozzle and an by the inherent simplicity of solid electromechanical actuator. rocket motors. Only the Attitude Control Castor IVATM solid rocket motors I System, which will be fueled prior to may be strapped on to the LL V family of assembly on the pad, uses liquid vehicles to increase payload to orbit propellant. The payload (spacecraft) capability. will be mated to the payload adapter and I release mechanism in a clean room. Attitude Control System After payload checkout, the Fairing will The Orbital Assist Module(OAM) be installed and the resulting is located above the Orbus 21 D. It I Encapsulated Assembly will be moved contains the LLV Flight Electronics, and placed on top of the waiting solid Batteries, Telemetry, Inertial rocket motors on the pad. Final Measurement Unit, and the Attitude preparation and check out on the pad Control System (ACS). The ACS is a I includes installation of batteries, liquid monopropellant hydrazine destruct arming devices and final propulsion system which has 10 rocket electrical checks. The time from engine assemblies for pitch, roll, yaw I completion of payload receipt inspection control and velocity addition to correct for any errors induced during solid I I I 3 I rocket motor boost flight. The ACS may of altitude and 0.020 inclination of the be configured with 2, 4 or 6 hydrazine desired orbit. tanks so that the propellant load may be I tailored to the specific mission. Fuel Payload Accommodations load with the modular tanks is 260, Of specific interest to the 520, and 780 pounds. The ACS is potential user of a new launch vehicle manufactured by the Rocket Research are the payload accommodations I Company which provides four 50 Ibf features. These accommodations consist axial thrusters for velocity addition and of the fairings, payload adapters and six 25 Ibf thrusters for pitch, yaw and release mechanisms, on pad air­ I roll control. The hydrazine tanks which conditioning, T -0 umbilical and support are spun aluminum with a graphite equipment interfaces and electrical composite overwrap, are pressurized interfaces provid i ng te Ie metry with gaseous nitrogen to 460 psig and monitoring capability before and during I contains a elastomeric bladder for flight. The LLV offers some unique positive expulsion of the fuel. The ACS features in this area. system may use up to 98% of the fuel on The LLV family has been designed I board at an average specific impulse of with the underlying theme of reducing 220 seconds. The total impulse system cost. This is especially reflected available is 57,200 Ibf sec. All in the selection of available fairings I components of the ACS have a proven range from a 92 inch diameter fairing flight history which minimizes intended primarily for the LLV1 to the development cost and program risk. 141 inch fairing for the larger vehicles. The fairings are of a two piece clam I Guidance, Navigation & Control shell design. Zip tube, a patented clean The heart of the LLV Guidance, system, cuts the base and two halves Navigation and Control (GN&C) system allowing springs to provide the I is the Litton LN-100L Inertial separation forces. The increased volume Navigation Unit (INU). This is a strap of these fairings, which are large for down system which employs a this class of launch vehicle, simplifies nondithered Litton Zero-lock™ Laser the satellite packaging requirements for I Gyro(ZLGTM) and A-4 accelerometer deployables such as solar arrays and technologies together with a antennas. There is, however, a sophisticated 24-state Kalman filter.
Load more

Recommended publications
  • Contingency Shuttle Crew Support (Cscs)/Rescue Flight Resource Book
    CSCS/Rescue Flight Resource Book JSC-62900 CONTINGENCY SHUTTLE CREW SUPPORT (CSCS)/RESCUE FLIGHT RESOURCE BOOK OVERVIEW 1 Mission Operations CSCS 2 Directorate RESCUE 3 FLIGHT DA8/Flight Director Office Final July 12, 2005 National Aeronautics and Space Administration Lyndon B. Johnson Space Center Houston, Texas FINAL 07/12/05 2-i Verify this is the correct version before using. CSCS/Rescue Flight Resource Book JSC-62900 CONTINGENCY SHUTTLE CREW SUPPORT (CSCS)/RESCUE FLIGHT RESOURCE BOOK FINAL JULY 12, 2005 PREFACE This document, dated May 24, 2005, is the Basic version of the Contingency Shuttle Crew Support (CSCS)/Rescue Flight Resource Book. It is requested that any organization having comments, questions, or suggestions concerning this document should contact DA8/Book Manager, Flight Director Office, Building 4 North, Room 3039. This is a limited distribution and controlled document and is not to be reproduced without the written approval of the Chief, Flight Director Office, mail code DA8, Lyndon B. Johnson Space Center, Houston, TX 77058. FINAL 07/12/05 2-ii Verify this is the correct version before using. CSCS/Rescue Flight Resource Book JSC-62900 1.0 - OVERVIEW Section 1.0 is the overview of the entire Contingency Shuttle Crew Support (CSCS)/Rescue Flight Resource Book. FINAL 07/12/05 2-iii Verify this is the correct version before using. CSCS/Rescue Flight Resource Book JSC-62900 This page intentionally blank. FINAL 07/12/05 2-iv Verify this is the correct version before using. CSCS/Rescue Flight Resource Book JSC-62900 2.0 - CONTINGENCY SHUTTLE CREW SUPPORT (CSCS) 2.1 Procedures Overview.......................................................................................................2-1 2.1.1 ................................................................................
    [Show full text]
  • 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:............................................................................................
    [Show full text]
  • 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.
    [Show full text]
  • 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)
    [Show full text]
  • 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.
    [Show full text]
  • 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.
    [Show full text]
  • 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.
    [Show full text]
  • The Annual Compendium of Commercial Space Transportation: 2017
    Federal Aviation Administration The Annual Compendium of Commercial Space Transportation: 2017 January 2017 Annual Compendium of Commercial Space Transportation: 2017 i Contents About the FAA Office of Commercial Space Transportation The Federal Aviation Administration’s Office of Commercial Space Transportation (FAA AST) licenses and regulates U.S. commercial space launch and reentry activity, as well as the operation of non-federal launch and reentry sites, as authorized by Executive Order 12465 and Title 51 United States Code, Subtitle V, Chapter 509 (formerly the Commercial Space Launch Act). FAA AST’s mission is to ensure public health and safety and the safety of property while protecting the national security and foreign policy interests of the United States during commercial launch and reentry operations. In addition, FAA AST is directed to encourage, facilitate, and promote commercial space launches and reentries. Additional information concerning commercial space transportation can be found on FAA AST’s website: http://www.faa.gov/go/ast Cover art: Phil Smith, The Tauri Group (2017) Publication produced for FAA AST by The Tauri Group under contract. NOTICE Use of trade names or names of manufacturers in this document does not constitute an official endorsement of such products or manufacturers, either expressed or implied, by the Federal Aviation Administration. ii Annual Compendium of Commercial Space Transportation: 2017 GENERAL CONTENTS Executive Summary 1 Introduction 5 Launch Vehicles 9 Launch and Reentry Sites 21 Payloads 35 2016 Launch Events 39 2017 Annual Commercial Space Transportation Forecast 45 Space Transportation Law and Policy 83 Appendices 89 Orbital Launch Vehicle Fact Sheets 100 iii Contents DETAILED CONTENTS EXECUTIVE SUMMARY .
    [Show full text]
  • 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.
    [Show full text]
  • 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.
    [Show full text]
  • Upper Stages Using Liquid Propulsion and Metallized - Propellants
    NASA NASA-TP-3191 19920007933 Technical Paper 3191 February 1992 Upper Stages Using Liquid Propulsion and Metallized - Propellants NASA NASA Technical Paper 3191 1992 Upper Stages Using Liquid Propulsion and Metallized Propellants Bryan A. Palaszewski Lewis Research Center Cleveland, Ohio NASA National Aeronauticsand Space Administration Office of Management Scientific and Technical InformationProgram Summary (Isp)may be required. Also, because of the limits of the capa- bilityof the IUS and potentiallylimitedavailabilityof the Titan Metallized propellants are liquid propellants with a metal IV/Centaur G-Prime for NASA missions, alternativesto these additive suspended in a gelled fuel. Typically, aluminum par- stages should be considered. ticles are the metal additive. These propellants increase the The largest available stage for the STS is the Inertial Upper density and/or the specific impulse of the propulsion system. Stage (IUS, ref. 1). It can deliver a 2268-kg (5000-Ibm)pay- Usingmetallized propellantsfor volume-and mass-constrained loadto geosynchronousEarth orbit (GEO).However, for plane- upper stages can deliver modest increases inperformance for tary missions, the IUS is limited to low-energy missions. The low Earth orbit to geosynchronous Earth orbit (LEO-GEO) t Galileo mission to Jupiter (ref. 2) was launched on an IUS. and other Earth-orbital transfer missions. However, using Using the Space Transportation System (STS)/IUS, its flight metallized propellants for planetary missionscan deliver great time will be 6.5 yr. With a high-performance cryogenic upper reductions in flight time with a single-stage, upper-stage stage, the flight time could havebeen reduced to 1.5 yr. There system, is not, however, any cryogenic upper stage available that is Tradeoff studies comparing metallized propellant stage compatible with the STS (refs.
    [Show full text]
  • Access to the Outer Solar System
    Interstellar Probe Study Webinar Series Access to the Outer Solar System Presenters • Michael Paul • Program Manager, Interstellar Probe Study, Johns Hopkins APL • Robert Stough • Payload Utilization Manager for NASA’s SLS Rocket, NASA Marshall Space Flight Center 12:05 PM EDT Thursday, 9 July 2020 Interstellar Probe Study Website http://interstellarprobe.jhuapl.edu NOW: The Heliosphere and the Local Interstellar Medium Our Habitable Astrosphere Sol G2V Main Sequence Star 24 km/s Habitable Mira BZ Camelopardalis LL Orionis IRC+10216 Zeta Ophiuchi Interstellar Probe Study 9 July 2020 2 Voyager – The Accidental Interstellar Explorers Uncovering a New Regime of Space Physics Global Topology Cosmic Ray Shielding Unexpected Field Direction Force Balance Not Understood Required Hydrogen Wall Measured (Voyager) Interstellar Probe Study 9 July 2020 3 Opportunities Across Disciplines Modest Cross-Divisional Contributions with High Return Extra-Galactic Background Light Dwarf Planets and KBOs Early galaxy and star formation Solar system formation Today Arrokoth Big Bang Pluto 13.7 Gya First Stars & Galaxies Circum-Solar Dust Disk ~13Gya Imprint of solar system evolution Sol 4.6 Ga HL-Tau 1 Ma! Poppe+2019 Interstellar Probe Study 9 July 2020 4 Earth-Jupiter-Saturn Sequences • Point Solutions indicated per year (capped at C3 = 312.15km2/s2) C3 Speed Dest. 2037 2 2 Year Date (km /s ) CA_J (rJ) CA_S (rS) (AU/yr) (Lon, Lat) 2036 17 Sept 182.66 1.05 2.0 5.954 (247,0) 2038 2037 15 Oct 312.15 1.05 2.0 7.985 (230,0) 2038 14 Nov 312.15 1.05 2.0 7.563 (241,0.1) 2036 2039 2039 15 Nov 274.65 1.05 2.0 5.055 (256,0.3) Interstellar Probe Study 9 July 2020 5 SPACE LAUNCH SYSTEM INTERSTELLAR PROBE Robert Stough SLS Spacecraft/Payload Integration & Evolution (SPIE) July 9, 2020 0760 SLS EVOLVABILITY FOUNDATION FOR A GENERATION OF DEEP SPACE EXPLORATION 322 ft.
    [Show full text]