Electric Propulsion and Plasma Dynamics Laboratory

Total Page:16

File Type:pdf, Size:1020Kb

Electric Propulsion and Plasma Dynamics Laboratory P1: ZBU Final Pages Qu: 00, 00, 00, 00 Encyclopedia of Physical Science and Technology EN005C-201 June 15, 2001 20:23 Electric Propulsion Robert G. Jahn Edgar Y. Choueiri Princeton University I. Conceptual Organization and History of the Field II. Electrothermal Propulsion III. Electrostatic Propulsion IV. Electromagnetic Propulsion V. Systems Considerations VI. Applications GLOSSARY Resistojet Device that heats a propellant stream by pass- ing it through a resistively heated chamber before Arcjet Device that heats a propellant stream by passing the propellant is expanded through a downstream a high-current electrical arc through it, before the pro- nozzle. pellant is expanded through a downstream nozzle. Thrust Unbalanced internal force exerted on a rocket Hall effect Conduction of electric current perpendicular during expulsion of its propellant mass. to an applied electric field in a superimposed magnetic field. Inductive thruster Device that heats a propellant stream THE SCIENCE AND TECHNOLOGY of electric by means of an inductive discharge before the propel- propulsion (EP) encompass a broad variety of strate- lant is expanded through a downstream nozzle. gies for achieving very high exhaust velocities in order Ion thruster Device that accelerates propellant ions by to reduce the total propellant burden and corresponding an electrostatic field. launch mass of present and future space transportation Magnetoplasmadynamic thruster Device that acceler- systems. These techniques group broadly into three cat- ates a propellant plasma by an internal or external mag- egories: electrothermal propulsion, wherein the propel- netic field acting on an internal arc current. lant is electrically heated, then expanded thermodynami- Plasma Heavily ionized state of matter, usually gaseous, cally through a nozzle; electrostatic propulsion, wherein composed of ions, electrons, and neutral atoms or ionized propellant particles are accelerated through molecules, that has sufficient electrical conductivity to an electric field; and electromagnetic propulsion, wherein carry substantial current and to react to electric and current driven through a propellant plasma interacts magnetic body forces. with an internal or external magnetic field to provide a Encyclopedia of Physical Science and Technology, Third Edition, Volume 5 Copyright C 2002 by Academic Press. All rights of reproduction in any form reserved. 125 P1: ZBU Final Pages Encyclopedia of Physical Science and Technology EN005C-201 June 15, 2001 20:23 126 Electric Propulsion stream-wise body force. Such systems can produce a range υe and logarithmically on the amount of propellant mass of exhaust velocities and payload mass fractions an order expended: of magnitude higher than that of the most advanced chem- mo ical rockets, which can thereby enable or substantially υ = υe ln , (2) m enhance many attractive space missions. The attainable f thrust densities (thrust per unit exhaust area) of these sys- where mo and m f are the total spacecraft mass at the start tems are much lower, however, which predicates longer and completion of the acceleration period. Conversely, the flight times and more complex mission trajectories. In ad- deliverable mass fraction, m f /mo, is a negative exponen- dition, these systems require space-borne electric power tial in the scalar ratio υ/υe: supplies of low specific mass and high reliability, inter- m f = eυ/υe . (3) faced with suitable power processing equipment. Opti- m mization of EP systems thus involves multidimensional o trade-offs among mission objectives, propellant and power Inclusion of significant gravitational or drag forces on plant mass, trip time, internal and external environmental the flight of the spacecraft adds appropriate terms to Eq. (1) factors, and overall system reliability. An enduring inter- and considerably complicates its integration, but it is still national program of research and development of viable possible to retain relation (3), provided that υ is now electric thrusters has been in progress for several decades, regarded as a more generalized “characteristic velocity and over the past few years this has led to the increas- increment,” indicative of the energetic difficulty of the ing use of a number of EP systems on commercial and particular mission or maneuver. However represented, the governmental spacecraft. Meanwhile, yet more advanced salient point is simply that if the spacecraft is to deliver a EP concepts have matured to high credibility for future significant portion of its initial mass to its destination, the mission applications. rocket exhaust speed must be comparable to this charac- teristic velocity increment. Clearly, for missions of large υ, the burden of thrust generation must shift from high rates of ejection of propellant mass to high relative exhaust I. CONCEPTUAL ORGANIZATION velocities. Unfortunately, conventional chemical rockets, AND HISTORY OF THE FIELD whether liquid or solid, monopropellant or bipropellant, are fundamentally limited by their available combustion A. Motivation reaction energies and heat transfer tolerances to exhaust The stimulus for development of electrically driven speeds of a few thousand meters per second, whereas space propulsion systems is nothing less fundamen- many attractive space missions entail characteristic ve- tal than Newton’s laws of dynamics. Since a rocket- locity increments at least an order of magnitude higher. propelled spacecraft in free flight derives its only Thus, some fundamentally difierent concept for the accel- acceleration from discharge of propellant mass, its equa- eration of propellant mass that circumvents the intrinsic tion of motion follows directly from conservation of limitations of chemical thermodynamic expansion is re- the total momentum of the spacecraft and its exhaust quired. Into this breech step the family of electric propul- stream: sion possibilities. mυú = mú υ , (1) e B. Conceptual Subdivision where m is the mass of the spacecraft at any given time, So that propellant exhaust speeds in the range above υú its acceleration vector, υe the velocity vector of the ex- 10,000 m/sec desirable for interplanetary flight and other haust jet relative to the spacecraft, and mú the rate of change high-energy missions can be obtained, processes basically of spacecraft mass due to propellant-mass expulsion. The different from nozzled expansion of a chemically reacting product mú υe is called the thrust of the rocket, T, and for flow must be invoked. More intense forms of propellant most purposes can be treated as if it were an external force heating may be employed, provided that the walls of the applied to the spacecraft. Its integral over any given thrust- rocket chamber and nozzle are protected from excessive ing time is usually termed the impulse, I, and the ratio of heat transfer. Alternatively, the thermal expansion route the magnitude of T to the rate of expulsion of propellant may be bypassed completely by direct application of suit- in units of sea-level weight, mgú o, has historically been la- able body forces to accelerate the propellant stream. Either beled the specific impulse, Is = υe/go.Ifυeis constant of these options is most reasonably accomplished by elec- over a given period of thrust, the spacecraft achieves an trical means, which constitute the technology of electric increment in its velocity, υ, which depends linearly on propulsion. P1: ZBU Final Pages Encyclopedia of Physical Science and Technology EN005C-201 June 15, 2001 20:23 Electric Propulsion 127 Historically, conceptually, and pragmatically, this field necessity for impeccable reliability of both components has tended to subdivide into three categories: of the system over long periods of unattended operation in the space environment. 1. Electrothermal propulsion, wherein the propellant is heated by some electrical process, then expanded C. History of Effort through a suitable nozzle 2. Electrostatic propulsion, wherein the propellant is The attractiveness of EP for a broad variety of space trans- accelerated by direct application of electrostatic portation applications was recognized by the patriarch of forces to ionized particles modern rocketry, Robert H. Goddard, as early as 1906. His 3. Electromagnetic propulsion, wherein the propellant is Russian counterpart, Konstantin Tsiolkovskiy, proposed accelerated under the combined action of electric and similar concepts in 1911, as did the German Hermann magnetic fields Oberth in his classic book on spaceflight in 1929 and the British team of Shepherd and Cleaver in 1949. But the first Over their periods of development, each of these ap- systematic and tutorial assessment of EP systems should proaches has spawned its own array of technical special- be attributed to Ernst Stuhlinger, whose book Ion Propul- ties and subspecialties, its own balance sheet of advantages sion for Space Flight nicely summarizes his seminal stud- and limitations, and its own cadres of proponents and de- ies of the 1950s. tractors, but in serious assessment, each has validly qual- The rapid acceleration of the U.S. space ambitions in ified for particular niches of application, many of which the 1960s drove with it the first coordinated research and do not seriously overlap. Throughout the history of EP de- development programs explicitly addressing EP technol- velopment, the original subdivision of the field into elec- ogy. In its earliest phase, this efiort drew heavily on reser- trothermal, electrostatic, and electromagnetic systems has voirs of past experience
Recommended publications
  • Call for M5 Missions
    ESA UNCLASSIFIED - For Official Use M5 Call - Technical Annex Prepared by SCI-F Reference ESA-SCI-F-ESTEC-TN-2016-002 Issue 1 Revision 0 Date of Issue 25/04/2016 Status Issued Document Type Distribution ESA UNCLASSIFIED - For Official Use Table of contents: 1 Introduction .......................................................................................................................... 3 1.1 Scope of document ................................................................................................................................................................ 3 1.2 Reference documents .......................................................................................................................................................... 3 1.3 List of acronyms ..................................................................................................................................................................... 3 2 General Guidelines ................................................................................................................ 6 3 Analysis of some potential mission profiles ........................................................................... 7 3.1 Introduction ............................................................................................................................................................................. 7 3.2 Current European launchers ...........................................................................................................................................
    [Show full text]
  • Launch and Deployment Analysis for a Small, MEO, Technology Demonstration Satellite
    46th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2008-1131 7 – 10 January 20006, Reno, Nevada Launch and Deployment Analysis for a Small, MEO, Technology Demonstration Satellite Stephen A. Whitmore* and Tyson K. Smith† Utah State University, Logan, UT, 84322-4130 A trade study investigating the economics, mass budget, and concept of operations for delivery of a small technology-demonstration satellite to a medium-altitude earth orbit is presented. The mission requires payload deployment at a 19,000 km orbit altitude and an inclination of 55o. Because the payload is a technology demonstrator and not part of an operational mission, launch and deployment costs are a paramount consideration. The payload includes classified technologies; consequently a USA licensed launch system is mandated. A preliminary trade analysis is performed where all available options for FAA-licensed US launch systems are considered. The preliminary trade study selects the Orbital Sciences Minotaur V launch vehicle, derived from the decommissioned Peacekeeper missile system, as the most favorable option for payload delivery. To meet mission objectives the Minotaur V configuration is modified, replacing the baseline 5th stage ATK-37FM motor with the significantly smaller ATK Star 27. The proposed design change enables payload delivery to the required orbit without using a 6th stage kick motor. End-to-end mass budgets are calculated, and a concept of operations is presented. Monte-Carlo simulations are used to characterize the expected accuracy of the final orbit.
    [Show full text]
  • Propulsion Options for the Global Precipitation Measurement Core Satellite
    PROPULSION OPTIONS FOR THE GLOBAL PRECIPITATION MEASUREMENT CORE SATELLITE Eric H. Cardiff*, Gary T. Davist, and David C. FoltaS NASA Goddard Space Flight Center Greenbelt, MD 2077 1 Abstract This study was conducted to evaluate several propulsion system options for the Global Precipitation Measurement (GPM) core satellite. Orbital simulations showed clear benefits for the scientific data to be obtained at a constant orbital altitude rather than with a decayheboost approach. An orbital analysis estimated the drag force on the satellite will be 1 to 12 mN during the five-year mission. Four electric propulsion systems were identified that are able to compensate for these drag forces and maintain a circular orbit. The four systems were the UK-lO/TS and the NASA 8 cm ion engines, and the ESA RMT and RITlO EVO radio-frequency ion engines. The mass, cost, and power requirements were examined for these four systems. The systems were also evaluated for the transfer time from the initial orbit of 400 x 650 km altitude orbit to a circular 400 km orbit. The transfer times were excessive, and as a consequence a “dual” system concept (with a hydrazine monopropellant system for the orbit transfer and electric propulsion for drag compensation) was examined. Clear mass benefits were obtained with the “dual” system, but cost remains an issue because of the larger power system required for the electric propulsion system. An electrodynamic tether was also evaluated in this trade study. Introduction The propulsion system is required to perform two The Global Precipitation Measurement (GPM) primary functions. The first is to transfer the mission will be launched in late 2008 to measure satellite from the launch insertion orbit to the final the amount and type of precipitation around the circular orbit.
    [Show full text]
  • The Propulsion of Sea Ships – in the Past, Present and Future –
    The Propulsion of Sea Ships – in the Past, Present and Future – (Speech by Bernd Röder on the occasion of the VHT General Meeting on 11.12.2008) To prepare for today’s topic, more specifically for the topic: ship propulsion of the future, I did what every reasonable person would have done in my situation if he should have a look into the future – I dug out our VHT crystal ball. As you know, our crystal ball is a reliable and cost-efficient resource which we have been using for a long time with great suc- cess. Among other things, we’ve been using it to provide you with the repair costs or their duration or the probable claims experience of a policy, etc. or to create short-term damage statistics, as well. So, I asked the crystal ball: What does ship propulsion of the future look like? Every future and all statistics lie in the crystal ball I must, however, admit that what I saw there was somewhat irritating, and it led me to only the one conclusion – namely, that we’re taking a look into the very distant future at a time in which humanity has not only used up all of the oil reserves but also the entire wind. This time, we’re not really going to get any further with the crystal ball. Ship propulsion of the future or 'back to the roots'? But, sometimes it helps to have a look into the past to be able to say something about the fu- ture. According to the principle, draw a line connecting the distant past to today and simply extend it into the future.
    [Show full text]
  • Rocket Nozzles: 75 Years of Research and Development
    Sådhanå Ó (2021) 46:76 Indian Academy of Sciences https://doi.org/10.1007/s12046-021-01584-6Sadhana(0123456789().,-volV)FT3](0123456789().,-volV) Rocket nozzles: 75 years of research and development SHIVANG KHARE1 and UJJWAL K SAHA2,* 1 Department of Energy and Process Engineering, Norwegian University of Science and Technology, 7491 Trondheim, Norway 2 Department of Mechanical Engineering, Indian Institute of Technology Guwahati, Guwahati 781039, India e-mail: [email protected]; [email protected] MS received 28 August 2020; revised 20 December 2020; accepted 28 January 2021 Abstract. The nozzle forms a large segment of the rocket engine structure, and as a whole, the performance of a rocket largely depends upon its aerodynamic design. The principal parameters in this context are the shape of the nozzle contour and the nozzle area expansion ratio. A careful shaping of the nozzle contour can lead to a high gain in its performance. As a consequence of intensive research, the design and the shape of rocket nozzles have undergone a series of development over the last several decades. The notable among them are conical, bell, plug, expansion-deflection and dual bell nozzles, besides the recently developed multi nozzle grid. However, to the best of authors’ knowledge, no article has reviewed the entire group of nozzles in a systematic and comprehensive manner. This paper aims to review and bring all such development in one single frame. The article mainly focuses on the aerodynamic aspects of all the rocket nozzles developed till date and summarizes the major findings covering their design, development, utilization, benefits and limitations.
    [Show full text]
  • Design of the Trajectory of a Tether Mission to Saturn
    DOI: 10.13009/EUCASS2019-510 8TH EUROPEAN CONFERENCE FOR AERONAUTICS AND AEROSPACE SCIENCES (EUCASS) DOI: Design of the trajectory of a tether mission to Saturn ⋆ Riccardo Calaon , Elena Fantino§, Roberto Flores ‡ and Jesús Peláez† ⋆Dept. of Industrial Engineering, University of Padova Via Gradenigio G/a -35131- Padova, Italy §Dept. of Aerospace Engineering, Khalifa University Abu Dhabi, Unites Arab Emirates ‡Inter. Center for Numerical Methods in Engineering Campus Norte UPC, Gran Capitán s/n, Barcelona, Spain †Space Dynamics Group, UPM School of Aeronautical and Space Engineering, Pz. Cardenal Cisneros 3, Madrid, Spain [email protected], [email protected], rfl[email protected], [email protected] †Corresponding author Abstract This paper faces the design of the trajectory of a tether mission to Saturn. To reduce the hyperbolic excess velocity at arrival to the planet, a gravity assist at Jupiter is included. The Earth-to-Jupiter portion of the transfer is unpropelled. The Jupiter-to-Saturn trajectory has two parts: a first coasting arc followed by a second arc where a low thrust engine is switched on. An optimization process provides the law of thrust control that minimizes the velocity relative to Saturn at arrival, thus facilitating the tethered orbit insertion. 1. Introduction The outer planets are of particular interest in terms of what they can reveal about the origin and evolution of our solar system. They are also local analogues for the many extra-solar planets that have been detected over the past twenty years. The study of these planets furthers our comprehension of our neighbourhood and provides the foundations to understand distant planetary systems.
    [Show full text]
  • An Assessment of Aerocapture and Applications to Future Missions
    Post-Exit Atmospheric Flight Cruise Approach An Assessment of Aerocapture and Applications to Future Missions February 13, 2016 National Aeronautics and Space Administration An Assessment of Aerocapture Jet Propulsion Laboratory California Institute of Technology Pasadena, California and Applications to Future Missions Jet Propulsion Laboratory, California Institute of Technology for Planetary Science Division Science Mission Directorate NASA Work Performed under the Planetary Science Program Support Task ©2016. All rights reserved. D-97058 February 13, 2016 Authors Thomas R. Spilker, Independent Consultant Mark Hofstadter Chester S. Borden, JPL/Caltech Jessie M. Kawata Mark Adler, JPL/Caltech Damon Landau Michelle M. Munk, LaRC Daniel T. Lyons Richard W. Powell, LaRC Kim R. Reh Robert D. Braun, GIT Randii R. Wessen Patricia M. Beauchamp, JPL/Caltech NASA Ames Research Center James A. Cutts, JPL/Caltech Parul Agrawal Paul F. Wercinski, ARC Helen H. Hwang and the A-Team Paul F. Wercinski NASA Langley Research Center F. McNeil Cheatwood A-Team Study Participants Jeffrey A. Herath Jet Propulsion Laboratory, Caltech Michelle M. Munk Mark Adler Richard W. Powell Nitin Arora Johnson Space Center Patricia M. Beauchamp Ronald R. Sostaric Chester S. Borden Independent Consultant James A. Cutts Thomas R. Spilker Gregory L. Davis Georgia Institute of Technology John O. Elliott Prof. Robert D. Braun – External Reviewer Jefferey L. Hall Engineering and Science Directorate JPL D-97058 Foreword Aerocapture has been proposed for several missions over the last couple of decades, and the technologies have matured over time. This study was initiated because the NASA Planetary Science Division (PSD) had not revisited Aerocapture technologies for about a decade and with the upcoming study to send a mission to Uranus/Neptune initiated by the PSD we needed to determine the status of the technologies and assess their readiness for such a mission.
    [Show full text]
  • Utilizing Deep Reinforcement Learning to Effect Autonomous Orbit Transfers and Intercepts Via Electromagnetic Propulsion
    Problem Utilizing Deep Reinforcement Learning to Effect Autonomous Results The growth in space-capable entities has caused a rapid rise in Orbit Transfers and Intercepts via Electromagnetic Propulsion ➢ Analysis the number of derelict satellites and space debris in orbit Gabriel Sutherland ([email protected]), Oregon State University, Corvallis, OR, USA ❖ The data suggests spacecraft highly capable of around Earth, which pose a significant navigation hazard. Frank Soboczenski ([email protected]), King’s College London, London, UK neutralizing/capturing small to intermediate debris Athanasios Vlontzos ([email protected]), Imperial College London, London, UK ❖ In simulations, spacecraft was able to capture and/or Objectives neutralize debris ranging in mass from 1g to 100g ❖ Satellites suffered damage in some Develop a system that is capable of autonomously neutralizing simulations multiple pieces of space debris in various orbits. Software & Simulations ❖ Simulations show that damaged segments of satellite were vaporized, which means Background that no additional debris was added to orbit ➢ Dangerous amounts of space debris in orbit, estimates vary ❖ Neural Network based on Deep Deterministic Policy ➢20,000+ tracked objects larger Gradient (DDPG) effective at low-thrust orbit transfers in than 10 cm diameter 2D Hohmann transfer problem ➢Est 500,000 objects larger ❖ Using ASTOS simulation software, mission-representative than 1 cm in diameter model of spacecraft and experimental propulsion system ➢Est 100 million objects ❖ Interfaced with DRL to train the Neural smaller than 1cm in diameter ➢ Orbital speeds of these objects Network with realistic data vary from 100+ kph to 28,100 kph ➢ Future Studies ➢ Spacecraft collisions due to space ❖ Train DDPG for 3 dimensional orbit transfer problems debris have been sparse so far Tracked spacecraft in Earth orbit.
    [Show full text]
  • VASIMR VX-200 Performance and Near-Term SEP Capability for Unmanned Mars Flight
    VASIMR VX-200 Performance and Near-term SEP Capability for Unmanned Mars Flight Future In-Space Operations Seminar January 19, 2011 presented by Tim Glover Director of Development [email protected] Ad Astra Rocket Company www.adastrarocket.com 1 Ad Astra Rocket Company Notes and Acronyms Notes: - solar array power values are for 1 AU - a number of publications on VASIMR R&D are available on the company’s website: http://www.adastrarocket.com Acronyms: SEP solar electric propulsion NTR nuclear thermal rocket VASIMR Variable Specific Impulse Magnetoplasma Rocket TMI Trans-Mars Insertion MOI Mars Orbit Insertion SOI sphere of influence IMLEO initial mass in low Earth orbit 2 Ad Astra Rocket Company Outline 1. VASIMR Prototype Performance 2. Simplified Earth‐Mars Trajectories 3. Chemical and NTR Hohmann Transfers 4. SEP: Initial Mass‐to‐Power Ratio and Payload Fraction 5. Near‐term SEP Mars Capabilities 6. Backup slides: • Propellant and transit time variation for actual orbits of Earth and Mars •Atlas V 551 performance curve 3 Ad Astra Rocket Company My Background • B.S., Physics, New Mexico State U. • M.S., Aerospace Engineering (Orbital Mechanics), UT-Austin • M.S., Physics, University of Pittsburgh • five years teaching high school physics, Ethical Culture Schools, New York • Ph.D., Applied Physics, Rice University, 2002 − thesis research at Johnson Space Center: designed and built plasma diagnostics to measure exhaust velocity in early VASIMR prototypes • Research Scientist, MEI Technologies 2003-2005 − continued experimental work on VASIMR up to 50 kW • Director of Development, Ad Astra Rocket Company 2005 – present − business development, external relations 4 Ad Astra Rocket Company VASIMR Operating Principles Superconducting Magnets typical path of an ion Helicon coupler (30 kW) through the rocket Ion cyclotron (170 kW) coupler i gas cold accelerated plasma plasma POWER 1.
    [Show full text]
  • Gravity-Assist Trajectories to Jupiter Using Nuclear Electric Propulsion
    AAS 05-398 Gravity-Assist Trajectories to Jupiter Using Nuclear Electric Propulsion ∗ ϒ Daniel W. Parcher ∗∗ and Jon A. Sims ϒϒ This paper examines optimal low-thrust gravity-assist trajectories to Jupiter using nuclear electric propulsion. Three different Venus-Earth Gravity Assist (VEGA) types are presented and compared to other gravity-assist trajectories using combinations of Earth, Venus, and Mars. Families of solutions for a given gravity-assist combination are differentiated by the approximate transfer resonance or number of heliocentric revolutions between flybys and by the flyby types. Trajectories that minimize initial injection energy by using low resonance transfers or additional heliocentric revolutions on the first leg of the trajectory offer the most delivered mass given sufficient flight time. Trajectory families that use only Earth gravity assists offer the most delivered mass at most flight times examined, and are available frequently with little variation in performance. However, at least one of the VEGA trajectory types is among the top performers at all of the flight times considered. INTRODUCTION The use of planetary gravity assists is a proven technique to improve the performance of interplanetary trajectories as exemplified by the Voyager, Galileo, and Cassini missions. Another proven technique for enhancing the performance of space missions is the use of highly efficient electric propulsion systems. Electric propulsion can be used to increase the mass delivered to the destination and/or reduce the trip time over typical chemical propulsion systems.1,2 This technology has been demonstrated on the Deep Space 1 mission 3 − part of NASA’s New Millennium Program to validate technologies which can lower the cost and risk and enhance the performance of future missions.
    [Show full text]
  • Propulsion and Flight Controls Integration for the Blended Wing Body Aircraft
    Cranfield University Naveed ur Rahman Propulsion and Flight Controls Integration for the Blended Wing Body Aircraft School of Engineering PhD Thesis Cranfield University Department of Aerospace Sciences School of Engineering PhD Thesis Academic Year 2008-09 Naveed ur Rahman Propulsion and Flight Controls Integration for the Blended Wing Body Aircraft Supervisor: Dr James F. Whidborne May 2009 c Cranfield University 2009. All rights reserved. No part of this publication may be reproduced without the written permission of the copyright owner. Abstract The Blended Wing Body (BWB) aircraft offers a number of aerodynamic perfor- mance advantages when compared with conventional configurations. However, while operating at low airspeeds with nominal static margins, the controls on the BWB aircraft begin to saturate and the dynamic performance gets sluggish. Augmenta- tion of aerodynamic controls with the propulsion system is therefore considered in this research. Two aspects were of interest, namely thrust vectoring (TVC) and flap blowing. An aerodynamic model for the BWB aircraft with blown flap effects was formulated using empirical and vortex lattice methods and then integrated with a three spool Trent 500 turbofan engine model. The objectives were to estimate the effect of vectored thrust and engine bleed on its performance and to ascertain the corresponding gains in aerodynamic control effectiveness. To enhance control effectiveness, both internally and external blown flaps were sim- ulated. For a full span internally blown flap (IBF) arrangement using IPC flow, the amount of bleed mass flow and consequently the achievable blowing coefficients are limited. For IBF, the pitch control effectiveness was shown to increase by 18% at low airspeeds.
    [Show full text]
  • Cave-73-02-Fullr.Pdf
    EDITORIAL Production Changes for Future Publication of the Journal of Cave and Karst Studies SCOTT ENGEL Production Editor The Journal of Cave and Karst Studies has experienced December 2011 issue, printed copies of the Journal will be budget shortfalls for the last several years for a multitude automatically distributed to paid subscribers, institutions, of reasons that include, but are not limited to, increased and only those NSS members with active Life and cost of paper, increased costs of shipping through the Sustaining level memberships. The remainder of the NSS United State Postal Service, increased submissions, and membership will be able to view the Journal electronically stagnant funding from the National Speleological Society online but will not automatically receive a printed copy. Full (NSS). The cost to produce the Journal has increased 5 to content of each issue of the Journal will be available for 20 percent per year for the last five years, yet the budget for viewing and downloading in PDF format at no cost from the the Journal has remained unchanged. To offset rising costs, Journal website www.caves.org/pub/journal. the Journal has implemented numerous changes over recent Anyone wishing to receive a printed copy of the Journal years to streamline the production and printing process. will be able to subscribe for an additional cost separate However, the increasing production costs, combined with from normal NSS dues. The cost and subscription process the increasing rate of good-quality submissions, has were still being determined at the time of this printing. resulted in the number of accepted manuscripts by the Once determined, the subscription information will be Journal growing faster than the acquisition of funding to posted on the Journal website.
    [Show full text]