Improving the Delivered Specific Impulse
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Aerospace Engine Data
AEROSPACE ENGINE DATA Data for some concrete aerospace engines and their craft ................................................................................. 1 Data on rocket-engine types and comparison with large turbofans ................................................................... 1 Data on some large airliner engines ................................................................................................................... 2 Data on other aircraft engines and manufacturers .......................................................................................... 3 In this Appendix common to Aircraft propulsion and Space propulsion, data for thrust, weight, and specific fuel consumption, are presented for some different types of engines (Table 1), with some values of specific impulse and exit speed (Table 2), a plot of Mach number and specific impulse characteristic of different engine types (Fig. 1), and detailed characteristics of some modern turbofan engines, used in large airplanes (Table 3). DATA FOR SOME CONCRETE AEROSPACE ENGINES AND THEIR CRAFT Table 1. Thrust to weight ratio (F/W), for engines and their crafts, at take-off*, specific fuel consumption (TSFC), and initial and final mass of craft (intermediate values appear in [kN] when forces, and in tonnes [t] when masses). Engine Engine TSFC Whole craft Whole craft Whole craft mass, type thrust/weight (g/s)/kN type thrust/weight mini/mfin Trent 900 350/63=5.5 15.5 A380 4×350/5600=0.25 560/330=1.8 cruise 90/63=1.4 cruise 4×90/5000=0.1 CFM56-5A 110/23=4.8 16 -
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. -
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 -
Enhancement of Volumetric Specific Impulse in HTPB/Ammonium Nitrate Mixed
Utah State University DigitalCommons@USU All Graduate Plan B and other Reports Graduate Studies 12-2016 Enhancement of Volumetric Specific Impulse in TPB/H Ammonium Nitrate Mixed Hybrid Rocket Systems Jacob Ward Forsyth Utah State University Follow this and additional works at: https://digitalcommons.usu.edu/gradreports Part of the Propulsion and Power Commons Recommended Citation Forsyth, Jacob Ward, "Enhancement of Volumetric Specific Impulse in TPB/AmmoniumH Nitrate Mixed Hybrid Rocket Systems" (2016). All Graduate Plan B and other Reports. 876. https://digitalcommons.usu.edu/gradreports/876 This Report is brought to you for free and open access by the Graduate Studies at DigitalCommons@USU. It has been accepted for inclusion in All Graduate Plan B and other Reports by an authorized administrator of DigitalCommons@USU. For more information, please contact [email protected]. ENHANCEMENT OF VOLUMETRIC SPECIFIC IMPULSE IN HTPB/AMMONIUM NITRATE MIXED HYBRID ROCKET SYSTEMS by Jacob W. Forsyth A report submitted in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE in Aerospace Engineering Approved: ______________________ ____________________ Stephen A. Whitmore Ph.D. David Geller Ph.D. Major Professor Committee Member ______________________ Rees Fullmer Ph.D. Committee Member UTAH STATE UNIVERSITY Logan, Utah 2016 ii Copyright © Jacob W. Forsyth 2016 All Rights Reserved iii ABSTRACT Enhancement of Volumetric Specific Impulse in HTPB/Ammonium Nitrate Mixed Hybrid Rocket Systems by Jacob W. Forsyth, Master of Science Utah State University, 2016 Major Professor: Dr. Stephen A. Whitmore Department: Mechanical and Aerospace Engineering Hybrid rocket systems are safer and have higher specific impulse than solid rockets. However, due to large oxidizer tanks and low regression rates, hybrid rockets have low volumetric efficiency and very long longitudinal profiles, which limit many of the applications for which hybrids can be used. -
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
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. -
An Investigation of the Performance Potential of A
AN INVESTIGATION OF THE PERFORMANCE POTENTIAL OF A LIQUID OXYGEN EXPANDER CYCLE ROCKET ENGINE by DYLAN THOMAS STAPP RICHARD D. BRANAM, COMMITTEE CHAIR SEMIH M. OLCMEN AJAY K. AGRAWAL A THESIS Submitted in partial fulfillment of the requirements for the degree of Master of Science in the Department of Aerospace Engineering and Mechanics in the Graduate School of The University of Alabama TUSCALOOSA, ALABAMA 2016 Copyright Dylan Thomas Stapp 2016 ALL RIGHTS RESERVED ABSTRACT This research effort sought to examine the performance potential of a dual-expander cycle liquid oxygen-hydrogen engine with a conventional bell nozzle geometry. The analysis was performed using the NASA Numerical Propulsion System Simulation (NPSS) software to develop a full steady-state model of the engine concept. Validation for the theoretical engine model was completed using the same methodology to build a steady-state model of an RL10A-3- 3A single expander cycle rocket engine with corroborating data from a similar modeling project performed at the NASA Glenn Research Center. Previous research performed at NASA and the Air Force Institute of Technology (AFIT) has identified the potential of dual-expander cycle technology to specifically improve the efficiency and capability of upper-stage liquid rocket engines. Dual-expander cycles also eliminate critical failure modes and design limitations present for single-expander cycle engines. This research seeks to identify potential LOX Expander Cycle (LEC) engine designs that exceed the performance of the current state of the art RL10B-2 engine flown on Centaur upper-stages. Results of this research found that the LEC engine concept achieved a 21.2% increase in engine thrust with a decrease in engine length and diameter of 52.0% and 15.8% respectively compared to the RL10B-2 engine. -
Propulsione Aeronautica 2020/2021 Francesco Barato
PROPULSIONE AERONAUTICA 2020/2021 FRANCESCO BARATO MATERIALE DI SUPPORTO FONDAMENTI DI PROPULSIONE AERONAUTICA Thrust 푇 = (푚̇ 푎 + 푚̇ 푓)푉푒 − 푚̇ 푎푉0 + (푝푒 − 푝푎)퐴푒 푇 ≈ 푚̇ 푎(푉푒 − 푉0) + (푝푒 − 푝푎)퐴푒 1 PROPULSIONE AERONAUTICA 2020/2021 FRANCESCO BARATO Ramjet P-270 Moskit (left), BrahMos (right) Turboramjet Pratt & Whitney J-58 turbo(ram)jet 2 PROPULSIONE AERONAUTICA 2020/2021 FRANCESCO BARATO Scramjet 3 PROPULSIONE AERONAUTICA 2020/2021 FRANCESCO BARATO Specific impulse 푇 푉푒 푇 푚̇ 푝 푉푒 − 푉0 퐼푠푝 = = [푠] 푟표푐푘푒푡푠 퐼푠푝 = = [푠] 푎푟 푏푟푒푎푡ℎ푛푔 푚̇ 푝푔0 푔0 푚̇ 푓푔0 푚̇ 푓 푔0 4 PROPULSIONE AERONAUTICA 2020/2021 FRANCESCO BARATO Propulsive efficiency Overall efficiency Overall efficiency with Mach number 5 PROPULSIONE AERONAUTICA 2020/2021 FRANCESCO BARATO Engine bypass ratios Bypass Engine Name Major applications ratio turbojet early jet aircraft, Concorde 0.0 SNECMA M88 Rafale 0.30 GE F404 F/A-18, T-50, F-117 0.34 PW F100 F-16, F-15 0.36 Eurojet EJ200 Typhoon 0.4 Klimov RD-33 MiG-29, Il-102 0.49 Saturn AL-31 Su-27, Su-30, J-10 0.59 Kuznetsov NK-144A Tu-144 0.6 PW JT8D DC-9, MD-80, 727, 737 Original 0.96 Soloviev D-20P Tu-124 1.0 Kuznetsov NK-321 Tu-160 1.4 GE Honda HF120 HondaJet 2.9 RR Tay Gulfstream IV, F70, F100 3.1 GE CF6-50 A300, DC-10-30,Lockheed C-5M Super Galaxy 4.26 PowerJet SaM146 SSJ 100 4.43 RR RB211-22B TriStar 4.8 PW PW4000-94 A300, A310, Boeing 767, Boeing 747-400 4.85 Progress D-436 Yak-42, Be-200, An-148 4.91 GE CF6-80C2 A300-600, Boeing 747-400, MD-11, A310 4.97-5.31 RR Trent 700 A330 5.0 PW JT9D Boeing 747, Boeing 767, A310, DC-10 5.0 6 PROPULSIONE -
A Tool for Preliminary Design of Rockets Aerospace Engineering
A Tool for Preliminary Design of Rockets Diogo Marques Gaspar Thesis to obtain the Master of Science Degree in Aerospace Engineering Supervisor : Professor Paulo Jorge Soares Gil Examination Committee Chairperson: Professor Fernando José Parracho Lau Supervisor: Professor Paulo Jorge Soares Gil Members of the Committee: Professor João Manuel Gonçalves de Sousa Oliveira July 2014 ii Dedicated to my Mother iii iv Acknowledgments To my supervisor Professor Paulo Gil for the opportunity to work on this interesting subject and for all his support and patience. To my family, in particular to my parents and brothers for all the support and affection since ever. To my friends: from IST for all the companionship in all this years and from Coimbra for the fellowship since I remember. To my teammates for all the victories and good moments. v vi Resumo A unica´ forma que a humanidade ate´ agora conseguiu encontrar para explorar o espac¸o e´ atraves´ do uso de rockets, vulgarmente conhecidos como foguetoes,˜ responsaveis´ por transportar cargas da Terra para o Espac¸o. O principal objectivo no design de rockets e´ diminuir o peso na descolagem e maximizar o payload ratio i.e. aumentar a capacidade de carga util´ ao seu alcance. A latitude e o local de lanc¸amento, a orbita´ desejada, as caracter´ısticas de propulsao˜ e estruturais sao˜ constrangimentos ao projecto do foguetao.˜ As trajectorias´ dos foguetoes˜ estao˜ permanentemente a ser optimizadas, devido a necessidade de aumento da carga util´ transportada e reduc¸ao˜ do combust´ıvel consumido. E´ um processo utilizado nas fases iniciais do design de uma missao,˜ que afecta partes cruciais do planeamento, desde a concepc¸ao˜ do ve´ıculo ate´ aos seus objectivos globais. -
Basic Analysis of a LOX/Methane Expander Bleed Engine
DOI: 10.13009/EUCASS2017-332 7TH EUROPEAN CONFERENCE FOR AERONAUTICS AND AEROSPACE SCIENCES (EUCASS) DOI: ADD DOINUMBER HERE Basic Analysis of a LOX/Methane Expander Bleed Engine ? ? ? Marco Leonardi , Francesco Nasuti † and Marcello Onofri ?Sapienza University of Rome Via Eudossiana 18, Rome, Italy [email protected] [email protected] [email protected] · · †Corresponding author Abstract As present trends in rocket engine development recommend overall simplicity and reliability as the main design driver, while preserving high performance, expander cycle engines based on the oxygen-methane pair have been considered as a possible upper stage option. A closed expander cycle is considered for Vega Evolution upper stage, while there are no studies published in the literature on methane-based expander bleed cycles. A basic cycle analysis is presented to evaluate the performance of an oxygen/methane ex- pander bleed cycle for an engine of 100 kN thrust class. Results show the feasibility of the system and its peculiarities with respect to the better known expander bleed cycle based on hydrogen. 1. Introduction The high chamber pressure required to achieve high specific impulse in liquid propellant rocket engines (LRE), has been efficiently obtained by pump-fed systems. Different solutions have been proposed since the beginning of space age and just a few of them has found its own field of application. In these systems the pumps are driven by gas turbines whose power comes from two possible sources: combustion or cooling system. The different needs for the specific applications (booster, sustainer or upper stage of different classes of rockets) led to classify pump-fed LRE systems in open and closed cycles, which differ because of turbine discharge pressure.14, 16 Closed cycles are those providing the best performance because the whole propellant mass flow rate is exploited in the main chamber. -
Space Shuttle Main Engine Orientation
BC98-04 Space Transportation System Training Data Space Shuttle Main Engine Orientation June 1998 Use this data for training purposes only Rocketdyne Propulsion & Power BOEING PROPRIETARY FORWARD This manual is the supporting handout material to a lecture presentation on the Space Shuttle Main Engine called the Abbreviated SSME Orientation Course. This course is a technically oriented discussion of the SSME, designed for personnel at any level who support SSME activities directly or indirectly. This manual is updated and improved as necessary by Betty McLaughlin. To request copies, or obtain information on classes, call Lori Circle at Rocketdyne (818) 586-2213 BOEING PROPRIETARY 1684-1a.ppt i BOEING PROPRIETARY TABLE OF CONTENT Acronyms and Abbreviations............................. v Low-Pressure Fuel Turbopump............................ 56 Shuttle Propulsion System................................. 2 HPOTP Pump Section............................................ 60 SSME Introduction............................................... 4 HPOTP Turbine Section......................................... 62 SSME Highlights................................................... 6 HPOTP Shaft Seals................................................. 64 Gimbal Bearing.................................................... 10 HPFTP Pump Section............................................ 68 Flexible Joints...................................................... 14 HPFTP Turbine Section......................................... 70 Powerhead........................................................... -
Increasing Launch Rate and Payload Capabilities
Aerial Launch Vehicles: Increasing Launch Rate and Payload Capabilities Yves Tscheuschner1 and Alec B. Devereaux2 University of Colorado, Boulder, Colorado Two of man's greatest achievements have been the first flight at Kitty Hawk and pushing into the final frontier that is space. Air launch systems aim to combine these great achievements into a revolutionary way to deliver satellites, cargo and eventually people into space. Launching from an aircraft has many advantages, including the ability to launch at any inclination and from above bad weather, which could delay ground launches. While many concepts for launching rockets from an airplane have been developed, very few have made it past the drawing board. Only the Pegasus and, to a lesser extent, SpaceShipOne have truly shown the feasibility of such a system. However, a recent push for rapid, small payload to orbit launches by the military and a general need for cheaper, heavy lift options are leading to an increasing interest in air launch methods. In order for the efficiency and flexibility of the system to be realized, however, additional funding and research are necessary. Nomenclature ALASA = airborne launch assist space access ALS = air launch system DARPA = defense advanced research project agency LEO = low Earth orbit LOX = liquid oxygen RP-1 = rocket propellant 1 MAKS = Russian air launch system MECO = main engine cutoff SS1 = space ship one SS2 = space ship two I. Introduction W hat typically comes to mind when considering launching people or satellites to space are the towering rocket poised on the launch pad. These images are engraved in the minds of both the public and scientists alike.