DOCSLIB.ORG

RL10A-3-3A Rocket Engine Modeling Project

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
RL10A-3-3A Rocket Engine Modeling Project https://ntrs.nasa.gov/search.jsp?R=19970010379 2020-06-16T02:41:52+00:00Z NASA Technical Memorandum 107318 RL10A-3-3A Rocket Engine Modeling Project Michael Binder NYMA, Inc. Brook Park, Ohio Thomas Tomsik and Joseph P. Veres Lewis Research Center Cleveland, Ohio January 1997 NationalAeronauticsand Space Administration TABLE OF CONTENTS Page ° Introduction ............................................................................................................................................... 1 2. Overview of the RL10A-3-3A Rocket Engine ............................................................................................ 2 2.1 Engine System Configuration and Operation ....................................................................... 2 2.2 Fuel Turbopump ..................................................................................................... 4 2.3 Oxygen (LOX) Pump ...................................................................................................................... 5 2.4 Regenerative Cooling Jacket ....................................................................................................... 5 2.5 Combustion Chamber and Nozzle ......................................................................................... 6 2.6 Valves, Ducts and Manifolds ............................................................................................................. 6 3. Project Organization and Goals ........................................................................................................................... 7 3.1 Turbomachinery Modeling Goals .................................................................................................... 7 3.2 Combustion and Heat Transfer Modeling Goals .......................................................... 8 3.3 Ducts, Manifolds, and Valves Modeling Goals ........................................................................ 8 3A System Modeling Goals .......................................................................................................... 8 4. Component Modeling Results ............................................................................................................................. 9 4.1 Turbomachinery Modeling Results ................................................................................................... 9 4.1.1 Verification of Pump Performance Test Data ............................................................. 9 4.1.2 Extension of Pump Maps for Start and Shutdown Conditions ................................... 9 4.1.2.1 Extension of Pump Maps to Start Conditions ............................................. 10 4.1.2.2 Extension of Pump Maps to Shut-Down Conditions .............................. 11 4.1.2.3 Effects of Density Changes on Pump Performance Models ................... 11 4.1.3 Verification of Fuel Turbine Performance Test Data .................................................. 12 4.1.4 Extension of Turbine Maps ........................................................................................ 12 4.2 Combustion and Heat Transfer Modeling Results ......................................................................... 13 4.2.1 Enhanced Combustion Gas Properties ........................................................................... 13 4.2.2 Cooling Jacket Heat Transfer Model ................................................................................... 13 4.2.3 Thrust Chamber Performance Calculations ........................................................................... 15 4.2_3.1 One-dimensional combustion model layout ....................................................... 15 4.2.3.2 Detailed modeling of the chamber injector ........................................................ 16 4.2.3.3 Nozzle performance models ........................................................................ 16 4.2.3.4 Two-phase flow through nozzle ........................................................... 17 4.2.4 Injector Heat-Transfer Calculations .................................................................... 17 4.3 Duct, Valve, and Manifold Modeling Results ............................................................ 18 4.3.1 Verification of Duct, Manifold Sizes ............................................................................. 18 4.3.2 Prediction of Fluid Frictional Resistances ...................................................................... 18 4.3.3 Modeling of Valve Actuator Mechanisms .................................................................... 19 4.3.4 Modeling of Critical Two-Phase Flow Through a Valve or Orifice ................................ 19 4.3.5 Model of Flow Through Venturi .............................................................................. 21 4.4 The New RL10A-3-3A System Model ............................................................................................ 22 4.4.1 Evaluation of Component Models/Integration With New System Model ....................... 21 4.4.2 Differences Between Start and Shutdown System Models .................................................. 22 5. ModelingUncertainties ..................................................................................... 22 5.1 Hardware Uncertainties .................................................................. .23 5.2 Valve Uncertainties .............................................................................................. 23 5.3 Uncertainty of Initial Conditions .......................................................................... 24 5.4 Uncertainty in Ignition Time ....................................................................................... 24 5.5 Interaction Between Uncertain Parameters ........................................................................... 25 6. Comparison of System Model Predictions With Test Data .................................................................... 25 6.1 Verification of Steady State Performance Predictions .................... 25 6-2 Verification for Start Transient Simulations .......................................................... 26 6.3 Verification of Shutdown Transient Simulations ....... 27 7. Discussion of Modeling Results ............................................................................................27 7.1 Discussion of Turbomachinery Investigation ................................................................................... 27 7-2 Discussion of Combustion and Heat Transfer Investigation ................................................ 28 7.3 Discussion of Duct, Manifold and Valve Investigation .................................................................. 29 7.4 Discussion of System Model Simulation Results .................................................................................. 29 8. Concluding Remarks .............................................................................................................................................. 29 9. Recommendations for Future Research ........................................................................................ 31 10. Def'mition of Terms ............................................................................................................. 33 11. Acknowledgments .......................................................................................................................... 34 12. Appendixes A RL10A-3-3A Engine System Model for ROCETS .................................................................................... 87 B----Component Modeling of RL10 Fuel and Oxidizer Pumps ............................................................... 89 C---Component Modeling of RL10 Fuel Turbine .................................................................... 109 D---Component Modeling of RL 10A-3-3A Cooling Jacket ................................................ 119 E---Component Modeling of RL10 Injector, Combustion Chamber, and Nozzle ................... 135 F---Component Modeling of RL10 Injector Heat Transfer ............................................................... 149 G---Component Modeling of RL10 Duct Flow ..................................................................... 157 H Modeling of Two-Phase (Liquid/Gas) Flow ................................................................... 163 I---Symbols ......................................................................................................................... 171 J--Glossary of Model Component Names ................................................................................................. 173 13. References ................................................................................................................................................... 175 14. Tables 2.2.1--Summary of Fuel Turbopump Characteristics .............................................................................. 4 2.3.1---Summary of LOX Pump Characteristics ............................................................................................ 5 2.4.1mSummary of Cooling Jacket Characteristics ............................................................................... 5 2.5.1--Summary of Combustion Chamber/Nozzle Characteristics ................................................ 6 2.6.1--Summary of Major Duct and Valve Characteristics ................................................................ 7 4-2.1mDescription of Combustion Property Tables for RL10 Model .....................................................
Load more

Recommended publications
  • Back to the the Future? 07> Probing the Kuiper Belt
    SpaceFlight A British Interplanetary Society publication Volume 62 No.7 July 2020 £5.25 SPACE PLANES: back to the the future? 07> Probing the Kuiper Belt 634089 The man behind the ISS 770038 Remembering Dr Fred Singer 9 CONTENTS Features 16 Multiple stations pledge We look at a critical assessment of the way science is conducted at the International Space Station and finds it wanting. 18 The man behind the ISS 16 The Editor reflects on the life of recently Letter from the Editor deceased Jim Beggs, the NASA Administrator for whom the building of the ISS was his We are particularly pleased this supreme achievement. month to have two features which cover the spectrum of 22 Why don’t we just wing it? astronautical activities. Nick Spall Nick Spall FBIS examines the balance between gives us his critical assessment of winged lifting vehicles and semi-ballistic both winged and blunt-body re-entry vehicles for human space capsules, arguing that the former have been flight and Alan Stern reports on his grossly overlooked. research at the very edge of the 26 Parallels with Apollo 18 connected solar system – the Kuiper Belt. David Baker looks beyond the initial return to the We think of the internet and Moon by astronauts and examines the plan for a how it helps us communicate and sustained presence on the lunar surface. stay in touch, especially in these times of difficulty. But the fact that 28 Probing further in the Kuiper Belt in less than a lifetime we have Alan Stern provides another update on the gone from a tiny bleeping ball in pioneering work of New Horizons.
    [Show full text]
  • Turbopump Design and Analysis Approach for Nuclear Thermal Rockets
    Turbopump Design and Analysis Approach for Nuclear Thermal Rockets Shu-cheng S. Chen1, Joseph P. Veres2, and James E. Fittje3 1NASA Glenn Research Center, Cleveland, Ohio 44135 2Chief, Compressor Branch, NASA Glenn Research Center, Cleveland, Ohio 44135 3Analex Corporation, 1100 Apollo Drive, Brook Park, Ohio 44142 1(216) 433-3585, [email protected] Abstract. A rocket propulsion system, whether it is a chemical rocket or a nuclear thermal rocket, is fairly complex in detail but rather simple in principle. Among all the interacting parts, three components stand out: they are pumps and turbines (turbopumps), and the thrust chamber. To obtain an understanding of the overall rocket propulsion system characteristics, one starts from analyzing the interactions among these three components. It is therefore of utmost importance to be able to satisfactorily characterize the turbopump, level by level, at all phases of a vehicle design cycle. Here at NASA Glenn Research Center, as the starting phase of a rocket engine design, specifically a Nuclear Thermal Rocket Engine design, we adopted the approach of using a high level system cycle analysis code (NESS) to obtain an initial analysis of the operational characteristics of a turbopump required in the propulsion system. A set of turbopump design codes (PumpDes and TurbDes) were then executed to obtain sizing and performance characteristics of the turbopump that were consistent with the mission requirements. A set of turbopump analyses codes (PUMPA and TURBA) were applied to obtain the full performance map for each of the turbopump components; a two dimensional layout of the turbopump based on these mean line analyses was also generated.
    [Show full text]
  • Development of Turbopump for LE-9 Engine
    Development of Turbopump for LE-9 Engine MIZUNO Tsutomu : P. E. Jp, Manager, Research & Engineering Development, Aero Engine, Space & Defense Business Area OGUCHI Hideo : Manager, Space Development Department, Aero Engine, Space & Defense Business Area NIIYAMA Kazuki : Ph. D., Manager, Space Development Department, Aero Engine, Space & Defense Business Area SHIMIYA Noriyuki : Space Development Department, Aero Engine, Space & Defense Business Area LE-9 is a new cryogenic booster engine with high performance, high reliability, and low cost, which is designed for H3 Rocket. It will be the first booster engine in the world with an expander bleed cycle. In the designing process, the performance requirements of the turbopump and other components can be concurrently evaluated by the mathematical model of the total engine system including evaluation with the simulated performance characteristic model of turbopump. This paper reports the design requirements of the LE-9 turbopump and their latest development status. Liquid oxygen 1. Introduction turbopump Liquid hydrogen The H3 rocket, intended to reduce cost and improve turbopump reliability with respect to the H-II A/B rockets currently in operation, is under development toward the launch of the first H3 test rocket in FY 2020. In rocket development, engine is an important factor determining reliability, cost, and performance, and as a new engine for the H3 rocket first stage, an LE-9 engine(1) is under development. A rocket engine uses a turbopump to raise the pressure of low-pressure propellant supplied from a tank, injects the pressurized propellant through an injector into a combustion chamber to combust it under high-temperature and high- pressure conditions.
    [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]
  • 6. Chemical-Nuclear Propulsion MAE 342 2016
    2/12/20 Chemical/Nuclear Propulsion Space System Design, MAE 342, Princeton University Robert Stengel • Thermal rockets • Performance parameters • Propellants and propellant storage Copyright 2016 by Robert Stengel. All rights reserved. For educational use only. http://www.princeton.edu/~stengel/MAE342.html 1 1 Chemical (Thermal) Rockets • Liquid/Gas Propellant –Monopropellant • Cold gas • Catalytic decomposition –Bipropellant • Separate oxidizer and fuel • Hypergolic (spontaneous) • Solid Propellant ignition –Mixed oxidizer and fuel • External ignition –External ignition • Storage –Burn to completion – Ambient temperature and pressure • Hybrid Propellant – Cryogenic –Liquid oxidizer, solid fuel – Pressurized tank –Throttlable –Throttlable –Start/stop cycling –Start/stop cycling 2 2 1 2/12/20 Cold Gas Thruster (used with inert gas) Moog Divert/Attitude Thruster and Valve 3 3 Monopropellant Hydrazine Thruster Aerojet Rocketdyne • Catalytic decomposition produces thrust • Reliable • Low performance • Toxic 4 4 2 2/12/20 Bi-Propellant Rocket Motor Thrust / Motor Weight ~ 70:1 5 5 Hypergolic, Storable Liquid- Propellant Thruster Titan 2 • Spontaneous combustion • Reliable • Corrosive, toxic 6 6 3 2/12/20 Pressure-Fed and Turbopump Engine Cycles Pressure-Fed Gas-Generator Rocket Rocket Cycle Cycle, with Nozzle Cooling 7 7 Staged Combustion Engine Cycles Staged Combustion Full-Flow Staged Rocket Cycle Combustion Rocket Cycle 8 8 4 2/12/20 German V-2 Rocket Motor, Fuel Injectors, and Turbopump 9 9 Combustion Chamber Injectors 10 10 5 2/12/20
    [Show full text]
  • Combustion Tap-Off Cycle
    College of Engineering Honors Program 12-10-2016 Combustion Tap-Off Cycle Nicole Shriver Embry-Riddle Aeronautical University, [email protected] Follow this and additional works at: https://commons.erau.edu/pr-honors-coe Part of the Aeronautical Vehicles Commons, Other Aerospace Engineering Commons, Propulsion and Power Commons, and the Space Vehicles Commons Scholarly Commons Citation Shriver, N. (2016). Combustion Tap-Off Cycle. , (). Retrieved from https://commons.erau.edu/pr-honors- coe/6 This Article is brought to you for free and open access by the Honors Program at Scholarly Commons. It has been accepted for inclusion in College of Engineering by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. Honors Directed Study: Combustion Tap-Off Cycle Date of Submission: December 10, 2016 by Nicole Shriver [email protected] Submitted to Dr. Michael Fabian Department of Aerospace Engineering College of Engineering In Partial Fulfillment Of the Requirements Of Honors Directed Study Fall 2016 1 1.0 INTRODUCTION The combustion tap-off cycle is also known as the “topping cycle” or “chamber bleed cycle.” It is an open liquid bipropellant cycle, usually of liquid hydrogen and liquid oxygen, that combines the fuel and oxidizer in the main combustion chamber. Gases from the edges of the combustion chamber are used to power the engine’s turbine and are expelled as exhaust. Figure 1.1 below shows a picture representation of the cycle. Figure 1.1: Combustion Tap-Off Cycle The combustion tap-off cycle is rather unconventional for rocket engines as it has only been put into practice with two engines.
    [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]
  • Validation of a Simplified Model for Liquid Propellant Rocket Engine Combustion Chamber Design
    IOP Conference Series: Materials Science and Engineering PAPER • OPEN ACCESS Validation of a simplified model for liquid propellant rocket engine combustion chamber design To cite this article: M Hegazy et al 2020 IOP Conf. Ser.: Mater. Sci. Eng. 973 012003 View the article online for updates and enhancements. This content was downloaded from IP address 170.106.33.14 on 25/09/2021 at 23:25 AMME-19 IOP Publishing IOP Conf. Series: Materials Science and Engineering 973 (2020) 012003 doi:10.1088/1757-899X/973/1/012003 Validation of a simplified model for liquid propellant rocket engine combustion chamber design M Hegazy1, H Belal2, A Makled3 and M A Al-Sanabawy4 1 M.Sc. Student, Rocket Department, Military Technical College, Egypt 2 Assistant Professor, Rocket Department, Military Technical College, Egypt 3 Associate Professor. Zagazig University, Egypt 4 Associate Professor. Rocket Department, Military Technical College, Egypt [email protected] Abstract. The combustion phenomena inside the thrust chamber of the liquid propellant rocket engine are very complicated because of different paths for elementary processes. In this paper, the characteristic length (L*) approach for the combustion chamber design will be discussed compared to the effective length (Leff) approach. First, both methods are introduced then applied for real LPRE. The effective length methodology is introduced starting from the basic model until developing the empirical equations that may be used in the design process. The classical procedure of L* was found to over-estimate the required cylindrical length in addition to the inherent shortcoming of not giving insight where to move to enhance the design.
    [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]
  • Cryogenic Technology & Rocket Engines
    ISSN (O): 2393-8609 International Journal of Aerospace and Mechanical Engineering Volume 2 – No.5, August 2015 Cryogenic Technology & Rocket Engines AKHIL GARG KARTIK JAKHU KISHAN SINGH ABHINAV B.Tech – Aerospace B.Tech – Aerospace B.Tech – Aerospace MAURYA Engg. Engg. Engg. B.Tech – Aerospace PUNJAB PUNJAB PUNJAB Engg. TECHNICAL TECHNICAL TECHNICAL PUNJAB UNIVERSITY, UNIVERSITY, UNIVERSITY, TECHNICAL JALANDHAR JALANDHAR JALANDHAR UNIVERSITY, akhilgarg.313@g kartik.lphawk@g kishansngh1996 JALANDHAR mail.com mail.com @gmail.com abhinavguru123 @gmail.com ABSTRACT 3.2 What is Cryogenic Rocket Engine? This paper is all about the rocket engine involving the use of A cryogenic rocket engine is a rocket engine that cryogenic technology at a cryogenic temperature (123K). This uses a cryogenic fuel or oxidizer, that is, its fuel or basically uses the liquid oxygen and liquid hydrogen as an oxidizer (or both) is gases liquefied and stored at oxidizer and fuel, which are very clean and non-pollutant very low temperatures. Notably, these engines were fuels compared to other hydrocarbon fuels like petrol, diesel, one of the main factors of the ultimate success in gasoline, LPG, CNG, etc., sometimes, liquid nitrogen is also reaching the Moon by the Saturn V rocket. used as an fuel. During World War II, when powerful rocket engines were first considered by the German, American and Keywords Soviet engineers independently, all discovered that Rocket engine, Cryogenic technology, Cryogenic temperature, rocket engines need high mass flow rate of both Liquid hydrogen and Oxygen. oxidizer and fuel to generate a sufficient thrust. At that time oxygen and low molecular weight 1.
    [Show full text]
  • The Modified Fuel Turbopump of 2Nd Stage Engine for H3 Launch Vehicle
    DOI: 10.13009/EUCASS2017-189 7TH EUROPEAN CONFERENCE FOR AERONAUTICS AND SPACE SCIENCES (EUCASS) The Modified Fuel Turbopump of 2nd stage engine for H3 launch vehicle Naoki Nagao*, Hideaki Nanri*, Koichi Okita*, Yohei Ishizu**, Shinnosuke Yabuki** and Shinichiro Kohno** *Japan Aerospace Exploration Agency 2-1-1 Sengen, Tsukuba-shi, Ibaraki 305-8505 **IHI Corporation 229 Tonogaya Mizuho-machi, Nishitama-gun, Tokyo 190-1297 Abstract The turbine of fuel turbopump of 2nd stage engine for H3 launch vehicle was changed from a partial admission type to a full admission type to meet the requirement of longer firing duration in a flight. The full admission type had an advantage of increasing the efficiency because of the uniform flow in circumference direction, on the other hand, had a disadvantage of decreasing the efficiency caused with the smaller height of turbine blade. The design of turbine was carefully modified and it was verified with experiments that the turbine had 10% higher performance than the previous design. 1. Introduction The next generation Japanese mainstay launch vehicle; H3 launch vehicle, has been developed since April 2014 to increase the launch capability of domestic launchers and to enhance the international competitiveness based on the “Basic Plan on Space Policy” [1]. As the result of system study on H3 launch vehicle, the second stage engine of H- IIA/B rocket, named LE-5B-2 is modified and adopted to H3 launch vehicle, because the performance of LE-5B-2 almost meets the requirement of the system and the high reliability of LE-5B series has been proved in the long-term operation.
    [Show full text]