DOCSLIB.ORG

Introduction to High Speed Airbreathing Propulsion Systems

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Introduction to High Speed Airbreathing Propulsion Systems Introduction to High Speed Airbreathing Propulsion Systems MAE 5540 - Propulsion Systems • Oxidizer … “liquid air” Have to take our own along So that the engine can breathe At high altitudes • How Much Oxidizer? … depending op chosen propellants … 4-9 times as much As the fuel we carry! MAE 5540 - Propulsion Systems Flight Without Oxidizer, I What if we could get enough oxygen from ambient air? X-43 airbreathing SCRAMjet engine What happens to our Isp? OK .. Lets loose the Oxidizer … and our Isp goes up by a factor Of 7! … but where do we the the Oxidizer for combustion? …. MAE 5540 - Propulsion Systems … from the atmosphere MAE 5540 - Propulsion Systems Flight Without Oxidizer, II Do Not Need Oxidizer here For a large portion of the Launch trajectory there is Plenty Enough air For combustion if you are going fast enough Need Oxidizer here MAE 5540 - Propulsion Systems Fuel Efficiencies of Various High Speed Propulsion Systems Isp Minimal turbo-machinery MAE 5540 - Propulsion Systems Operational Flight Envelope for Various Flight Vehicles MAE 5540 - Propulsion Systems Operations Payoffs for Airbreathing Launch • Decreased gross lift-off weight, resulting in smaller facilities and easier handling • Wider range of emergency landing sites for intact abort • Powered flyback/go-around & more margin at reduced power • Self-ferry & taxi capabilities • Greatly expanded launch windows (double or triple) • Rapid orbital rendezvous (up to three times faster than rockets) • Wider array of landing sites from orbit, with 2,000-mile cross range and increased range • Reduced sensitivity to weight growth MAE 5540 - Propulsion Systems Applications MAE 5540 - Propulsion Systems Amount of Rocket propellant required to launch a 1000 kg payload into orbit Titan IV Our total launch mass goes from space shuttle Rockets10,000 on kg/kg the Launch down Pad to are <1400 Mostly ALLkg/kg FUEL!! wow! Amount of Propellant required, kg "Venture-star" MAE 5540 - Propulsion Systems Specific Impulse, sec. Airbreathing Propulsion Basics Squeeze Bang Blow Suck MAE 5540 - Propulsion Systems Airbreathing (Jet) Propulsion Basics (cont’d)! • Ramjet Engine! • Simplest of all Jet Engine= (Credit Narayanan Komerath, Georgia Tech)! Concepts! • Compression and Power Extraction steps Performed Step 1-2! Step 3! Step 4! passively! • Requires High Speed inlet flow to operate! MAE 5540 - Propulsion Systems! Airbreathing (Jet) Propulsion Basics (cont’d)! • Ramjet Engine! Ramjets cannot produce thrust at zero airspeed; they cannot move an aircraft from a standstill. A ramjet powered vehicle, therefore, requires an assisted take-off like a rocket assist to accelerate it to a speed where it begins to produce thrust.! https://www.youtube.com/watch? v=1DG4f2umclk! MAE 5540 - Propulsion Systems! Airbreathing Propulsion Basics (cont’d)! Turbine • Turbojet/Turbo fan ! Combustor Nozzle engines Combustion! Compressor Cycle steps! Turbojet Compression and ! Power extraction! Inlet! steps use! Afterburner Turbo-machinery! Turbine To augment cycle! Compressor Afterburner Turbofan Inlet! Combustor Nozzle MAE 5540 - Propulsion Systems! Fan Airbreathing Propulsion Basics (cont’d)! turbojet. Simple turbine engine that produces all of its thrust from the exhaust from the turbine section. However, because all of the air is passing through the whole turbine, all of it must burn fuel.! MAE 5540 - Propulsion Systems! Airbreathing Propulsion Basics (cont’d)! Turbofan. Turbine primarily drives a fan at the front of the engine. Most engines drive the fan directly from the turbine. Part of the air enters the turbine section of the engine, and the rest is bypassed around the engine. In high-bypass engines, most of the air only goes through the fan and bypasses the rest of the engine and providing most of the thrust. ! A turbofan thus can be thought of as a turbojet being used to drive a ducted fan, with both of those contributing to the thrust. The ratio of the mass-flow of air bypassing the engine core compared to the mass-flow of air passing through the core is referred to as the bypass ratio. ! MAE 5540 - Propulsion Systems! Airbreathing Propulsion Basics (cont’d)! Turbofans have a number of benefits over simple turbojets. It is more energy efficient to accelerate a large air mass by a small amount, than a small air mass by a lot. A turbofan effectively increases the volume of air propelled with a corresponding decrease in average exhaust gas velocity. ! Larger, slower fan, positioned upwind of the compressor pushes air (relatively) slowly through both the core and the bypass. The air that makes it to the core is compressed and undergoes combustion. The air that makes it to the bypass is ducted around the core, and mixes with the exhaust gases when it exits the back of the engine. ! MAE 5540 - Propulsion Systems! Ideal Ramjet Cycle Analysis Step 1-2 Step 3 Step 4 Region Process Ideal Behavior Real Behavior A to 1(inlet) Isentropic flow P0,T0 constant P0 drop 1-2 (diffuser) Adiabatic P,T increase P0 drop Compression P0 drop 2-3 (burner) Heat Addition P0 constant, T0 P0 drop s Increase " !q% !s = $ ' > 0 # T & rev 3-4 (nozzle) Isentropic T0,P0 constant s Increase expansion MAE 5540 - Propulsion Systems !s > !srev T0 drop Ideal Ramjet Cycle Analysis T-s Diagram Step 1-2 Step 3 Step 4 T s MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet • Net Work Available --> work perform by system in step 4 minus work required for step 1-2 • Net heat input --> heat input during step 3 (combustion) - heat lost in exhaust plume (Credit Narayanan Komerath, Georgia Tech) Step 1-2 Step 3 Step 4 MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) • Ideal Cycle Efficiency(η) = (Net work output)/(Net heat input} Net Work • = (hA ! hB ) + (hC ! hD ) m Net Heat Input • = (hC ! hB ) m (Credit Narayanan Komerath, Georgia Tech) Step 1-2 Step 3 Step 4 MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) • Ideal Cycle Efficiency = (Net work output)/(Net heat input} " Net Work % $ • ' # m & (hA ( hB ) + (hC ( hD ) (hD ( hA ) ! = = = 1( " Net Heat Input % (hC ( hB ) (hC ( hB ) $ • ' # m & (Credit Narayanan Komerath, Georgia Tech) Step 1-2 Step 3 Step 4 MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) • Assume … Cpair ~ Cpproducts # C & p air %TD " TA ( C T " C T C p products D p air A $ p products ' ! = 1" ( ) = 1" C T " C T # C & ( p air C p products B ) p air % TC " TB ( C $ p products ' (Credit Narayanan Komerath, Georgia Tech) Step 1-2 Step 3 Step 4 MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) • For simplicity … let … Cpair ~ Cpproducts $ C ' p air &TD # TA ) C % p products ( (TD # TA ) ! " 1# = 1# $ C ' (TC # TB ) p air & TC # TB ) C % p products ( $ T T ' $ T T T ' D # A D # A B %& T T () %& T T T () = 1# C C = 1# C B C $ TB ' $ TB ' &1# ) &1# ) % TC ( % TC ( MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) • From C-->D flow is isentropic … ( )1 (Credit Narayanan Komerath, Georgia Tech) ( TD " PD % ! = $ ' TC # PC & Step 1-2 Step 3 Step 4 # ) "1 & # P & ) T T # T T T & % D " A B ( D " A B % $% P '( T T ( $% T T T '( $ C B C ' ! = 1" C B C = 1" # TB & # TB & %1" ( %1" ( $ TC ' $ TC ' MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) • Adiabatic compression across diffuser (Credit Narayanan Komerath, Georgia Tech) ! P $ ! P P0 P0 $ A = A ' A ' B = Step 1-2 Step 3 Step 4 # P & # P P P & " B % " 0A 0B B % ( ( ( )1 ( )1 ! $ ( )1 T ( )1 P ( ! P $ ( T ! 0B $ 0A T ! P $ 0B A * T = T * A = A # T & # T & P 0A 0B T # P & # P & " 0A % " B % 0B B " B % " 0A % MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) • Sub into efficiency (Credit Narayanan Komerath, Georgia Tech) equation Step 1-2 Step 3 Step 4 # ) "1 ) "1 ) "1 & # P & ) # P & ) # P & ) T % D " A 0B B ( % % ( % ( % ( ( $ PC ' $ PB ' $ P0 ' TC $% A '( ! = 1" # TB & %1" ( $ TC ' MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) (Credit Narayanan Komerath, Georgia Tech) ideal nozzle ! PA = PD ideal burner ! PB = PC Step 1-2 Step 3 Step 4 # ) "1 & # ) "1 & # P & ) T # P & ) %1" 0B B ( %T " 0B T ( ) "1 % ( ) "1 C % ( B % P0 TC ( % P0 ( ) % $ A ' ( ) % $ A ' ( # PA & $ ' # PA & $ ' ! = 1" % ( = 1" % ( $ PB ' # TB & $ PB ' (TC " TB ) %1" ( $ TC ' MAE 5540 - Propulsion Systems Thermodynamic Efficiency of Ideal Ramjet (cont’d) # ) "1 & (Credit Narayanan Komerath, Georgia Tech) # P & ) %T " 0B T ( ) "1 C % ( B % P0 ( ) % $ A ' ( # PA & $ ' ! = 1" % ( $ PB ' (TC " TB ) Step 1-2 Step 3 Step 4 i) As engine pressure ratio, PB/PA, goes up … η goes up ii) As combustor temperature difference TC-TB goes up … η goes up iii) As inlet total pressure ratio (P0B/P0A) goes down … (stagnation pressure loss goes up) … η goes down MAE 5540 - Propulsion Systems Why is total pressure recovery important? • Choked • ! +1 (! (1) Nozzle Throat m ! " 2 % p0 massflow * = $ ' A Rg # ! + 1& T0 • Thrust is proportional to engine massflow • • Thrust = me Ve ! mi Vi + ( peAe ! p" Ae ) MAE 5540 - Propulsion Systems Ideal Ramjet: Inlet and Diffuser • Take a Rocket motor• and “lop the top off” m fuel • normal • M > 1 m M < 1 m air shockwave air+fuel • Works Ok for subsonic, but for supersonic flow … can’t cram enough air down the tube • Result is a normal shock wave at the inlet lip MAE 5540 - Propulsion Systems Ideal Ramjet: Inlet and Diffuser (cont’d) • m fuel • normal • M > 1 m M < 1 m air shockwave air+fuel M ∞ M2 • Mechanical Energy is Dissipated into Heat • Huge Loss in Momentum
Load more

Recommended publications
  • 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
    [Show full text]
  • 2. Afterburners
    2. AFTERBURNERS 2.1 Introduction The simple gas turbine cycle can be designed to have good performance characteristics at a particular operating or design point. However, a particu­ lar engine does not have the capability of producing a good performance for large ranges of thrust, an inflexibility that can lead to problems when the flight program for a particular vehicle is considered. For example, many airplanes require a larger thrust during takeoff and acceleration than they do at a cruise condition. Thus, if the engine is sized for takeoff and has its design point at this condition, the engine will be too large at cruise. The vehicle performance will be penalized at cruise for the poor off-design point operation of the engine components and for the larger weight of the engine. Similar problems arise when supersonic cruise vehicles are considered. The afterburning gas turbine cycle was an early attempt to avoid some of these problems. Afterburners or augmentation devices were first added to aircraft gas turbine engines to increase their thrust during takeoff or brief periods of acceleration and supersonic flight. The devices make use of the fact that, in a gas turbine engine, the maximum gas temperature at the turbine inlet is limited by structural considerations to values less than half the adiabatic flame temperature at the stoichiometric fuel-air ratio. As a result, the gas leaving the turbine contains most of its original concentration of oxygen. This oxygen can be burned with additional fuel in a secondary combustion chamber located downstream of the turbine where temperature constraints are relaxed.
    [Show full text]
  • A Supersonic Fan Equipped Variable Cycle Engine for a Mach 2.7 Supersonic Transport
    University of New Hampshire University of New Hampshire Scholars' Repository Applied Engineering and Sciences Scholarship Applied Engineering and Sciences 8-22-1985 A Supersonic Fan Equipped Variable Cycle Engine for a Mach 2.7 Supersonic Transport Theodore S. Tavares University of New Hampshire, Manchester, [email protected] Follow this and additional works at: https://scholars.unh.edu/unhmcis_facpub Recommended Citation Tavares, T.S., “A Supersonic Fan Equipped Variable Cycle Engine for a Mach 2.7 Supersonic Transport,” NASA-CR-177141, 1986. This Report is brought to you for free and open access by the Applied Engineering and Sciences at University of New Hampshire Scholars' Repository. It has been accepted for inclusion in Applied Engineering and Sciences Scholarship by an authorized administrator of University of New Hampshire Scholars' Repository. For more information, please contact [email protected]. https://ntrs.nasa.gov/search.jsp?R=19860019474 2018-07-25T19:53:49+00:00Z ^4/>*>?/ GAS TURBINE LABORATORY DEPARTMENT OF AERONAUTICS AND ASTRONAUTICS MASSACHUSETTS INSTITUTE OF TECHNOLOGY CAMBRIDGE, MA 02139 A FINAL REPORT ON NASA GRANT NAG-3-697 entitled A SUPERSONIC FAN EQUIPPED VARIABLE CYCLE ENGINE FOR A MACH 2.7 SUPERSONIC TRANSPORT by T. S. Tavares prepared for NASA Lewis Research Center Cleveland, OH 44135 (NASA-CB-177141) A SDPEBSCNIC FAN EQUIPPED N86-28946 VARIABLE CYCLE ENGINE.JOB A MACH 2.? SDPEESONIC TBANSPOBT Final Report (Massachusetts Inst. of Tech.) 107 p Unclas CSCL 21E G3/07 43461 August 22, 1985 A SUPERSONIC FAN EQUIPPED VARIABLE CYCLE ENGINE FOR A HACK 2.7 SUPERSONIC TRANSPORT by Theodore Sean Tavares A SUPERSONIC FAN EQUIPPED VARIABLE CYCLE ENGINE FOR A MACH 2.7 SUPERSONIC TRANSPORT by THEODORE SEAN TAVARES ABSTRACT A design stud/ was carried out to evaluate the concept of a variable cycle turbofan engine with an axially supersonic fan stage as powerplant for a Mach 2.7 supersonic transport.
    [Show full text]
  • 04 Propulsion
    Aircraft Design Lecture 2: Aircraft Propulsion G. Dimitriadis and O. Léonard APRI0004-­1, Aerospace Design Project, Lecture 4 1 Introduction • A large variety of propulsion methods have been used from the very start of the aerospace era: – No propulsion (balloons, gliders) – Muscle (mostly failed) – Steam power (mostly failed) – Piston engines and propellers – Rocket engines – Jet engines – Pulse jet engines – Ramjet – Scramjet APRI0004-­1, Aerospace Design Project, Lecture 4 2 Gliding flight • People have been gliding from the-­ mid 18th century. The Albatross II by Jean Marie Le Bris-­ 1849 Otto Lillienthal , 1895 APRI0004-­1, Aerospace Design Project, Lecture 4 3 Human-­powered flight • Early attempts were less than successful but better results were obtained from the 1960s onwards. Gerhardt Cycleplane (1923) MIT Daedalus (1988) APRI0004-­1, Aerospace Design Project, Lecture 4 4 Steam powered aircraft • Mostly dirigibles, unpiloted flying models and early aircraft Clément Ader Avion III (two 30hp steam engines, 1897) Giffard dirigible (1852) APRI0004-­1, Aerospace Design Project, Lecture 4 5 Engine requirements • A good aircraft engine is characterized by: – Enough power to fulfill the mission • Take-­off, climb, cruise etc. – Low weight • High weight increases the necessary lift and therefore the drag. – High efficiency • Low efficiency increases the amount fuel required and therefore the weight and therefore the drag. – High reliability – Ease of maintenance APRI0004-­1, Aerospace Design Project, Lecture 4 6 Piston engines • Wright Flyer: One engine driving two counter-­ rotating propellers (one port one starboard) via chains. – Four in-­line cylinders – Power: 12hp – Weight: 77 kg APRI0004-­1, Aerospace Design Project, Lecture 4 7 Piston engine development • During the first half of the 20th century there was considerable development of piston engines.
    [Show full text]
  • An Online Performance Calculator for Ramjet, Turbojet, Turbofan, and Turboprop Engines
    An Online Performance Calculator for Ramjet, Turbojet, Turbofan, and Turboprop Engines Evan P. Warner∗ Virginia Tech, Blacksburg, VA, 24061, USA This paper documents the theory, equations, and assumptions behind an HTML-based online performance calculator for ramjet, turbojet, turbofan, and turboprop engines. Based on inputs of design parameters, flight con- ditions, and engine parameters, the specific thrust (()), thrust specific fuel consumption ()(), propulsive efficiency ([ ?), thermal efficiency ([Cℎ), and overall efficiency ([0) will be output for the engine of interest. Several sample cases are analyzed to verify the functionality of the calculator. These sample cases are derived from engines associated with Virginia Tech, which include: Pratt & Whitney’s PT6A-20 and JT15D-1, Honeywell’s TFE731-2B, and the Rolls Royce Trent 1000. With the help of this calculator, students will now be able to verify their air-breathing jet engine performance value calculations. This calculator is available here. I. Nomenclature "0 = Ambient Mach Number 2 ?1 = Specific Heat of the Burner W0 = Ambient Specific Heat Ratio 2 ?C = Specific Heat of the Turbine W3 = Diffuser Specific Heat Ratio 2 ? 5 = Specific Heat of the Fan W2 = Compressor Specific Heat Ratio )0 = Ambient Static Temperature W1 = Burner Specific Heat Ratio ?0 = Ambient Static Pressure WC = Turbine Specific Heat Ratio ?4 = Exit Static Pressure W= = Nozzle Specific Heat Ratio '08A = Gas Constant for Air W 5 = Fan Specific Heat Ratio '?A>3 = Gas Constant for the Products of Fuel/Air Mixture W= 5 = Fan Nozzle
    [Show full text]
  • Modelling and Simulation of the Revolutionary Turbine Accelerator
    Announcement “This final year project was an exam. Commentary made during the presentation is not taken into account.” "Deze eindverhandeling was een examen. De tijdens de verdediging geformuleerde opmerkingen werden niet opgenomen". Preface Here I would like to thank everyone who made a contribution to my thesis. Also to all the people I forgot or do not call by name below. First of all I would like to thank my both VKI promoters. Guillermo Paniagua and Victor Fernández Villacé for their guidance, theoretical support and the help finding the exact information I was searching for. Also thanks to my KHBO promoter, Wim Vanparys who also gave me support, in addition to his clear view on the project and future considerations. Second I would like to say thank you to JeanFrançois Herbiet for his EcosimPro help, in which Victor also made his contribution, and for the information about the conical inlet and its implementation. Without the IT help of Olivier Jadot there would be no EcosimPro at the VKI, on my computer and it definitely would not be accessible at home. Also the people of the VKI library deserve a great thank you for delivering the requested papers as soon as possible which were essential to understand the working principle of the RTA. Next, a thank you to my VKI-colleagues, Piet Van den Ecker, Marylène Andre and Maarten De Moor for the great time and discussions about the occurring problems. Most of the time they helped me, without them knowing it, determining and resolving the problems. Of course to all the others as well, to keep up the spirit in the little room were we spent our time.
    [Show full text]
  • ITU Beengine for a Next Generation Trainer
    ITU BeEngine for a Next Generation Trainer AIAA 2015 - 2016 Undergraduate Engine Design Competition Team ITU BeEngine Proposal Advisor : Assoc. Prof. Dr. Onur TUNÇER Team ITU BeEngine Olcay SARI Orçun BULAT ABSTRACT The aim of this ITU BeEngine Design Project is to design a conceptual new engine to replace GE J-85-5A engine of T-38 supersonic jet trainer with better fuel consumption, performance and weight properties. New designed BeEngine is capable of providing the required thrust values with an improved fuel consumption values. On the other hand, since the purpose of T-38 is to train pilots for 5th generation fighter aircrafts, better performance properties are included into new engine. With the involvement of latest materials properties, such as TiAl alloys and CMC (Ceramic Matrix Composites), engine weight is decreased significantly for a similar but a little smaller size. Under the mixed flow low bypass turbofan category, new engine pushes the limits of overall pressure ratio of 16 through 8 compressor stages with a slightly higher bypass ratio of 1.2. Counter rotating 2-spool new engine is aimed to provide optimum values for prepared mission profile for trainer aircrafts while providing extended range of ≈1800nmi. With all ultimate technology limits, each spool is supported by a single stage turbine. Additionally, BeEngine is quite flexible on the afterburner operations thanks to its higher bypass ratio. Apart from achieving the required afterburner thrust values, it is also possible to increase maximum thrust limit in order to compete with other ultimate supersonic jet trainers in market. Furthermore, with new designed convergent-divergent nozzle with vectoring for allowing extra maneuverability for training purposes.
    [Show full text]
  • Innovative Airbreathing Propulsion Concepts for Access to Space
    r I I 2 00/Z20/03 NASA/TM-2001-210564 S 5lf'<f5 LCfj ( I I ~ Innovative Airbreathing Propulsion Concepts for Access to Space Woodrow Whitlow, Jr., Richard A. Blech, and Isaiah M. Blankson Glenn Research Center, Cleveland, Ohio I i :::::J October 2001 The NASA STI Program Office ... in Profile Since its founding, NASA has been dedicated to • CONFERENCE PUBLICATION. Collected the advancement of aeronautics and space papers from scientific and technical science. The NASA Scientific and Technical conferences, symposia, seminars, or other Information (STI) Program Office plays a key part meetings sponsored or cosponsored by in helping NASA maintain this important role. NASA The NASA STI Program Office is operated by • SPECIAL PUBLICATION. Scientific, Langley Research Center, the Lead Center for technical, or historical information from NASA's scientific and technical information. The NASA programs, projects, and missions, NASA STI Program Office provides access to the often concerned with subjects having NASA STI Database, the largest collection of substantial public interest. aeronautical and space science STI in the world. The Program Office is also NASA's institutional • TECHNICAL TRANSLATION. English­ mechanism for disseminating the results of its language translations of foreign scientific research and development activities. These results and technical material pertinent to NASA's are published by NASA in the NASA STI Report mission. Series, which includes the following report types: Specialized services that complement the STI • TECHNICAL PUBLICATION. Reports of Program Office's diverse offerings include completed research or a major Significant creating custom thesauri, building customized phase of research that present the results of data bases, organizing and publishing research NASA programs and include extensive data results ..
    [Show full text]
  • Design and Control of a Variable Geometry Turbofan with an Independently Modulated Third Stream
    DESIGN AND CONTROL OF A VARIABLE GEOMETRY TURBOFAN WITH AN INDEPENDENTLY MODULATED THIRD STREAM DISSERTATION Presented in Partial Fulfillment of the Requirements for the Degree Doctor of Philosophy in the Graduate School of The Ohio State University By Ronald J. Simmons, M.S. * * * * * The Ohio State University 2009 Dissertation Committee: Professor Meyer Benzakein, Adviser Professor Richard Bodonyi Approved by Professor Jeffrey Bons Professor Jen-Ping Chen Adviser Professor Nicholas J. Kuprowicz Aerospace Engineering Graduate Program Distribution Statement A: Unlimited Distribution. Cleared for Public Release by AFRL/WS Public Affairs Case Number 88ABW-2009-1697 The views expressed in this article are those of the author and do not reflect the official policy or position of the United States Air Force, Department of Defense, or the U.S. Government. ABSTRACT Abstract Emerging 21st century military missions task engines to deliver the fuel efficiency of a high bypass turbofan while retaining the ability to produce the high specific thrust of a low bypass turbofan. This study explores the possibility of satisfying such competing demands by adding a second independently modulated bypass stream to the basic turbofan architecture. This third stream can be used for a variety of purposes including: providing a cool heat sink for dissipating aircraft heat loads, cooling turbine cooling air, and providing a readily available stream of constant pressure ratio air for lift augmentation. Furthermore, by modulating airflow to the second and third streams, it is possible to continuously match the engine‟s airflow demand to the inlet‟s airflow supply thereby reducing spillage and increasing propulsive efficiency. This research begins with a historical perspective of variable cycle engines and shows a logical progression to proposed architectures.
    [Show full text]
  • SAE1307- Aerospace Propulsion UNIT 1
    SCHOOL OF MECHANICAL ENGINEERING DEPARTMENT OF AERONAUTICAL ENGINEEING SAE1307- Aerospace Propulsion UNIT 1 1: Performance of turbojets, ramjets at high speeds – limitations. Need for supersonic combustion: The renewed interest in high-speed propulsion has led to increased activity in the development of the supersonic combustion ramjet engine for hypersonic flight applications. In this flight regime, the scramjet engine‟s specific thrust exceeds that of other propulsion systems. Use of air breathing propulsion systems like, scramjets from takeoff to the edges of the atmosphere has the potential to reduce costs of space launch considerably. The hypersonic flight regime is commonly considered to begin when velocities exceed Mach 6 Defence applications of scramjets in missiles is also very sought after due to the very short reaction times associated with high speed of the missile system Subsonic combustion, which technologically is easier to manage with the current knowledge, would be associated, in the hypersonic regime, with high stagnation temperatures that would lead to unacceptable dissociation levels, and hence an inability to materialize the energy rise expected through chemical reactions Combined cycle engines: No single-engine cycle exists that can efficiently cover the whole range of a flight from takeoff to orbit insertion; therefore combined cycles are of particular interest for the design of the scramjet cycle Limitations of Turbojets/Turbofans and Ramjets at High Speeds: Performance based differences between the different engine cycles are clearly illustrated in the fuel specific impulse, Specific impulse vs Mach number diagram shown. = Fig:1.1 Mach 3 flight regime The diagram shows that around Mach 3 flight regime the subsonic combustion ramjet becomes more efficient as a propulsive system in comparison with the turbine-based engines (turbojets of turbofans).
    [Show full text]
  • Introduction to Airbreathing Propulsion Systems
    Introduction to airbreathing propulsion systems Integrated project: AIAA aircraft design competition 1er master Aerospace 2020 Koen Hillewaert, Université de Liège Design of Turbomachines, A&M [email protected] Remy Princivalle, Principal Engineer, Safran Aero Boosters [email protected] 1. Balances, thrust and performance Drag / thrust definition 1. Balances, thrust and performance Airbreathing engines: acceleration of (clean) air mass flow ● Thrust = acceleration force of engine mass flow from flight vf to jet velocity vj ● Powers – Propulsive power → airplane acceleration : – Mechanical power → fluid acceleration : – Lost power ● Propulsive efficiency: ● Increasing propulsive efficiency for constant thrust – increase mass flow, decrease jet velocity – Ingest flow at speed lower than flight speed 1. Balances, thrust and performance Airbreathing engines: Boundary Layer Ingestion ● ● 1. Balances, thrust and performance Airbreathing engines: classical performance parameters ● Thermal energy ! = mf Δhf – Fuel mass flow rate: – Fuel to air ratio: – Fuel lower heating value: ● "verall efficiency propulsive power Pp versus thermal energy ! – Propulsive efficiency: – Thermal efficiency: ● Efficiency $ Thrust specific fuel consumption (TSFC) ● Compacity $ Specific thrust 2. Jet engines Generation of high” speed "et through e#pansion over no%%le Snecma/'' (lympus 593 Afterburning ,urbo"et -./ Leap -i$il ,urbofan Snecma /88 Afterburning Turbofan 2. Jet engines Generation of high” speed "et through e#pansion over no%%le ● Ingestion of ma air at flight speed in nacelle – Ram effect: increased total T and p due to relative Mach num!er Mf ● Increase total pressure and temperature – Mechanical : fan – Thermal : gas generator " #rayton – $fter!urning ● #(pansion over e(haust noz)le to ambient pressure – %ho&ed – $dapted 2. Jet engines -ore flow: Brayton thermodynamic cycle ● Thermodynamic cycle – $dia!atic compression ' → ( : – %om!ustion ( → ) : – $dia!atic e*pansion )+, : 2.
    [Show full text]
  • Air Breathing Propulsion.Pdf
    AERONAUTICAL ENGINEERING_MRCET (AUTONOMOUS INSTITUTION) AIR BREATHING PROPULSION (R18A2109) COURSE FILE II B. Tech II Semester (2019-2020) Prepared By Prof Vedantam Ravi Department of Aeronautical Engineering MALLA REDDY COLLEGE OF ENGINEERING & TECHNOLOGY (Autonomous Institution – UGC, Govt. of India) Affiliated to JNTU, Hyderabad, Approved by AICTE - Accredited by NBA & NAAC – ‘A’ Grade - ISO 9001:2015 Certified) Maisammaguda, Dhulapally (Post Via. Kompally), Secunderabad – 500100, Telangana State, India. 1 AERONAUTICAL ENGINEERING_MRCET (AUTONOMOUS INSTITUTION) MRCET VISION To become a model institution in the fields of Engineering, Technology and Management. To have a perfect synchronization of the ideologies of MRCET with challenging demands of International Pioneering Organizations. MRCET MISSION To establish a pedestal for the integral innovation, team spirit, originality and competence in the students, expose them to face the global challenges and become pioneers of Indian vision of modern society. MRCET QUALITY POLICY To pursue continual improvement of teaching learning process of Undergraduate and Post Graduate programs in Engineering & Management vigorously. To provide state of art infrastructure and expertise to impart the quality education. 1 AERONAUTICAL ENGINEERING_MRCET (AUTONOMOUS INSTITUTION) PROGRAM OUTCOMES Engineering Graduates will be able to: 1. Engineering knowledge: Apply the knowledge of mathematics, science, engineering fundamentals, and an engineering specialization to the solution of complex engineering problems. 2. Problem analysis: Identify, formulate, review research literature, and analyze complex engineering problems reaching substantiated conclusions using first principles of mathematics, natural sciences, and engineering sciences. 3. Design / development of solutions: Design solutions for complex engineering problems and design system components or processes that meet the specified needs with appropriate consideration for the public health and safety, and the cultural, societal, and environmental considerations.
    [Show full text]