The Development and Testing of Pulsed Detonation Engine Ground

Total Page:16

File Type:pdf, Size:1020Kb

The Development and Testing of Pulsed Detonation Engine Ground THE DEVELOPMENT AND TESTING OF PULSED DETONATION ENGINE GROUND DEMONSTRATORS by PHILIP KOSHY PANICKER Presented to the Faculty of the Graduate School of The University of Texas at Arlington in Partial Fulfillment of the Requirements for the Degree of DOCTOR OF PHILOSOPHY THE UNIVERSITY OF TEXAS AT ARLINGTON August 2008 Copyright © by Philip K. Panicker All Rights Reserved Psalm 53:1 (Please do not be one.) 2 Peter 3: 2-9 (Sign of the times.) Psalm 127:1 Except the LORD build the house, they labor in vain that build it: except the LORD keep the city, the watchman waketh but in vain. Mark 8:36, 37 For what shall it profit a man, if he shall gain the whole world, and lose his own soul? Or what shall a man give in exchange for his soul? Psalm 103:2-5 Bless the LORD, O my soul, and forget not all his benefits: Who forgiveth all thine iniquities; who healeth all thy diseases; Who redeemeth thy life from destruction; who crowneth thee with loving kindness and tender mercies; Who satisfieth thy mouth with good things; so that thy youth is renewed like the eagle's. Romans 10:13 For whosoever shall call upon the name of the Lord shall be saved. The Holy Bible, Authorized KJV ACKNOWLEDGEMENTS I would first like to thank Dr. Frank Lu for his long and continual support and encouragement throughout this educational endeavor. I am grateful for the freedom to pursue research in my own direction, mostly, and for the opportunity to work at the ARC. I also am thankful for the pep talks, the apt refocusing of my attention during wavering moments and for all the help with research. I am extremely indebted to Dr. Lu for his diligent editing and proofreading of my dissertation manuscript. I wish to thank Dr. Don Wilson, who helped me with my scholarships and assistantships and other aid during my study here. I am grateful to him for making it possible for me to travel to Canberra, Australia and Reno, NV, for the AIAA conferences. I am also very thankful for his advice on the research subjects and for reading and editing my dissertation. I want to thank Dr. George Emanuel for his advice and instructions during the brief time that he was on my committee. I would like to thank my committee members, Dr. Brian Dennis, Dr. Kamesh Subbarao and Dr. Albert Tong for their advice and feedback. I would like to thank Dr. J. Craig Dutton, who during his time here, encouraged me in my research work. I am also grateful to Dr. Wen S. Chan, who as graduate advisor, arranged for my teaching assistantships. In addition, I am obliged to Dr. S.M. You for granting me the opportunity to teach the MAE2381 class for two semesters. I would like to thank the administrative staff of the MAE department, including Louella Carpenter, Janet Gober, Sally Thompson, Barbara Sanderson and Donna Woodhead, for their timely management of all bureaucratic procedures and for all their help in maintaining the smooth flow of operations without a hitch. Connie Niece will always be remembered for her friendly words and for her help with travel arrangements. I would like to thank Rodney Duke, Engineering iv Technician at the ARC, who converted all the articles within my drawings and without into tangible material and for all other help in setting up our research apparatus. I wish to thank Jiun-Ming (Jimmy) Li, who was my research partner during his one-year stint here at UTA and ARC, during which time we drew up designs, performed experiments and wrote papers together. I want to thank Dr. Daniel T.S. New, who got me started on the research on PDEs and also took on the bigger role in writing papers of our research. I am also grateful to Kin Fong Chui and Kim Seng Lim, who joined us here for research on PDEs. I want to thank Naveen Kumar Reddy and Harsha Shivashankar, who were co-researchers on the early PDEs. I want to thank Rafaela Bellini, for her love, friendship and support for which I am truly indebted. I am grateful to her for wielding her great skills with MS Word for the formatting and typesetting of this manuscript and all other help she extended to me during my days writing this work. I am very grateful to Albert Ortiz, my colleague and co-researcher, for all his invaluable help and partnership in performing the experiments and setting up equipment and instruments, and analyzing data; and as co-GTA of MAE2381. I want to thank my other colleagues here at ARC, Eric Braun, Richard Mitchell, Adam Pierce and Prashaanth Ravindran, who helped with research work and in running the PDEs. Along with the other inhabitants of the ARC, they made life here interesting and fun. I want to thank my father P. Koshy Panicker, my mother Deenamma for their love, prayers and continued support for all my education. I also want to thank my brother, John, who supported me financially and spiritually in my education. I want to thank my sister, Anu Koshy and brother-in-law Binu John, for their love. I want to thank the Kuruvilla family for all their love, help and prayers during my life here in DFW. I want to thank all who helped me along the way and could not be included by name here in this short list. Above all, I want to thank my Lord and Savior Jesus Christ, who gave me life and blessed my life abundantly. May all glory be to God. July 23, 2008 v ABSTRACT THE DEVELOPMENT AND TESTING OF PULSED DETONATION ENGINE GROUND DEMONSTRATORS Philip K. Panicker, PhD. The University of Texas at Arlington, 2008 Supervising Professor: Frank K. Lu The successful implementation of a PDE running on fuel and air mixtures will require fast-acting fuel-air injection and mixing techniques, detonation initiation techniques such as DDT enhancing devices or a pre-detonator, an effective ignition system that can sustain repeated firing at high rates and a fast and capable, closed-loop control system. The control system requires high-speed transducers for real-time monitoring of the PDE and the detection of the detonation wave speed. It is widely accepted that the detonation properties predicted by C-J detonation relations are fairly accurate in comparison to experimental values. The post- detonation flow properties can also be expressed as a function of wave speed or Mach number. Therefore, the PDE control system can use C-J relations to predict the post-detonation flow properties based on measured initial conditions and compare the values with those obtained from using the wave speed. The controller can then vary the initial conditions within the vi combustor for the subsequent cycle, by modulating the frequency and duty cycle of the valves, to obtain optimum air and fuel flow rates, as well as modulate the energy and timing of the ignition to achieve the required detonation properties. Five different PDE ground demonstrators were designed, built and tested to study a number of the required sub-systems. This work presents a review of all the systems that were tested, along with suggestions for their improvement. The PDE setups, ranged from a compact PDE with a 19 mm (3/4 in.) i.d., to two 25 mm (1 in.) i.d. setups, to a 101 mm (4 in.) i.d. dual-stage PDE setup with a pre-detonator. Propane-oxygen mixtures were used in the smaller PDEs. In the dual-stage PDE, propane- oxygen was used in the pre-detonator, while propane-air mixtures were used in the main combustor. Both rotary valves and solenoid valve injectors were studied. The rotary valves setups were tested at 10 Hz, while the solenoid valves were tested at up to 30 Hz on a 25 mm i.d. PDE. The dual-stage PDE was run at both 1 Hz and 10 Hz using solenoid valves. The two types of valves have their drawbacks and advantages which are discussed, along with ways to enhance their functionality. Rotary valves with stepper motor drives are recommended to be used for air flow control, while an array of solenoid injectors may be used for liquid or gaseous fuel injection. Various DDT enhancing devices were tested, including Shchelkin spirals (with varying thicknesses, lengths and pitches), grooved sleeves and converging-diverging nozzles. The Shchelkin spirals are found to be the most effective of all, at blockage ratios in the region of 50 to 55 %. To improve the durability of Shchelkin spirals, it is recommended that they be grooved into the inside of tubes or inserted as replaceable sleeves. Orifice plates with high blockage ratios, in the region of 50 to 80 %, are also recommended due to their simple and rugged design. All these devices along with the PDE combustor will require a strong cooling system to prevent damage from the extreme detonation temperatures. High energy (HE) and low energy (LE) ignition systems were tested and compared along with various designs of igniters and automotive spark plugs. It is concluded that while HE ignition may help unsensitized fuel-air mixtures to achieve detonations faster than LE systems, the former have severe drawbacks. The HE igniters get damaged quickly, and require large and heavy power vii supplies. While the HE ignition is able to reduce ignition delay in a propane-oxygen pre- detonator, it did not show a significant improvement in bringing about DDT in the main combustor using propane-air mixtures. The compact pre-detonator design with a gradual area change transitioning to a larger combustor is found to be effective for detonation initiation, but the pre-detonator concept is recommended for high-speed applications only, since higher speeds requires more sensitive, easily detonable fuels that have short ignition delays and DDT run-up distances.
Recommended publications
  • Considerations and Simulations About Pulse Detonation Engine
    MATEC Web of Conferences 290, 04009 (2019) https://doi.org/10.1051/matecconf /2019290 04009 MSE 2019 Considerations and simulations about Pulse Detonation Engine Vasile Prisacariu1,*, Constantin Rotaru1, and Mihai Leonida Niculescu2 1Henri Coandă Air Force Academy, Aviation Department, Brașov, Romania 2INCAS, Aerodynamic Department, București, Romania Abstract. PDE propulsion can work from a subsonic regime to hypersonic regimes; this type of engine can have higher thermodynamic efficiency compared to other turbojet or turbofan engines due to the removal of rotating construction elements (compressors and turbines) that can reduce the mass and total cost of propulsion system. The PDE experimental researches focused on both the geometric configuration and the thermo- gas-dynamic flow aspects to prevent uncontrolled self-ignition. This article presents a series of numerical simulations on the functioning of PDE with hydrogen at supersonic regimens. Acronyms and symbols PDE pulse detonation engine AB PDE after burner pulse detonation engine DDT deflagration-to-detonation transition PDRE pulse detonation rocket engine ZDN Zel'dovich-von Neumann-Doering Mx, My Mach numbers p pressure v volume T temperature ρ density u1, u2 velocity h1, h2 enthalpy s1, s2 (S) entropy Vo,V∞ speed 1 INTRODUCTION Detonation is an effective means of burning a fuel mixture and transforming chemical energy into mechanical energy. However, this concept of running the propulsion systems involves difficulties both in rapidly achieving the fuel mixture at high speeds and in initiating and sustaining detonation in a controlled manner. PDE differs from conventional propulsion systems in two main aspects: it generates an intermittent pulse and produces high pressure increase in the combustion chamber, which also represents a major advantage of a PDE, see Figure 1.
    [Show full text]
  • The SKYLON Spaceplane
    The SKYLON Spaceplane Borg K.⇤ and Matula E.⇤ University of Colorado, Boulder, CO, 80309, USA This report outlines the major technical aspects of the SKYLON spaceplane as a final project for the ASEN 5053 class. The SKYLON spaceplane is designed as a single stage to orbit vehicle capable of lifting 15 mT to LEO from a 5.5 km runway and returning to land at the same location. It is powered by a unique engine design that combines an air- breathing and rocket mode into a single engine. This is achieved through the use of a novel lightweight heat exchanger that has been demonstrated on a reduced scale. The program has received funding from the UK government and ESA to build a full scale prototype of the engine as it’s next step. The project is technically feasible but will need to overcome some manufacturing issues and high start-up costs. This report is not intended for publication or commercial use. Nomenclature SSTO Single Stage To Orbit REL Reaction Engines Ltd UK United Kingdom LEO Low Earth Orbit SABRE Synergetic Air-Breathing Rocket Engine SOMA SKYLON Orbital Maneuvering Assembly HOTOL Horizontal Take-O↵and Landing NASP National Aerospace Program GT OW Gross Take-O↵Weight MECO Main Engine Cut-O↵ LACE Liquid Air Cooled Engine RCS Reaction Control System MLI Multi-Layer Insulation mT Tonne I. Introduction The SKYLON spaceplane is a single stage to orbit concept vehicle being developed by Reaction Engines Ltd in the United Kingdom. It is designed to take o↵and land on a runway delivering 15 mT of payload into LEO, in the current D-1 configuration.
    [Show full text]
  • Study and Numerical Simulation of Unconventional Engine Technology
    STUDY AND NUMERICAL SIMULATION OF UNCONVENTIONAL ENGINE TECHNOLOGY by ANJALI SHEKHAR B.E Aeronautical Engineering VTU, Karnataka, 2013 A thesis submitted in partial fulfillment of the requirements for the degree of Master of Science, Aerospace Engineering, College of Engineering and Applied Science, University of Cincinnati, Ohio 2018 Thesis Committee: Chair: Ephraim Gutmark, Ph.D. Member: Shaaban Abdallah, Ph.D. Member: Mark Turner, Sc.D. An Abstract of Study and Numerical Simulation of Unconventional Engine Technology by Anjali Shekhar Submitted to the Graduate Faculty as partial fulfillment of the requirements for the Master of Science Degree in Aerospace Engineering University of Cincinnati December 2018 The aim of this thesis is to understand the working of two unconventional aircraft propul- sion systems and to setup a two-dimensional transient simulation to analyze its operational mechanism. The air traffic has nearly increased by about 40% in past three decades and calls for alternative propulsion techniques to replace or support the current traditional propulsion methodology. In the light of current demand, the thesis draws motivation from renewed inter- est in two non-conventional propulsion techniques designed in the past and had not been given due importance due to various flaws/drawbacks associated. The thesis emphasizes on the work- ing of Von Ohains thermal compression engine and pulsejet combustors. Computational Fluid Dynamics is used in current study as it offers very high flexibility and can be modified easily to incorporate the required changes. Thermal Compression engine is a design suggested by Von Ohain in 1948. The engine works on the principle of pressure rise caused inside the engine which completely depends on the temperature of working fluid and independent of rotations per minute.
    [Show full text]
  • Modelling a Hypersonic Single Expansion Ramp Nozzle of a Hypersonic Aircraft Through Parametric Studies
    energies Article Modelling a Hypersonic Single Expansion Ramp Nozzle of a Hypersonic Aircraft through Parametric Studies Andrew Ridgway, Ashish Alex Sam * and Apostolos Pesyridis College of Engineering, Design and Physical Sciences, Brunel University London, London UB8 3PH, UK; [email protected] (A.R.); [email protected] (A.P.) * Correspondence: [email protected]; Tel.: +44-1895-267-901 Received: 26 September 2018; Accepted: 7 December 2018; Published: 10 December 2018 Abstract: This paper aims to contribute to developing a potential combined cycle air-breathing engine integrated into an aircraft design, capable of performing flight profiles on a commercial scale. This study specifically focuses on the single expansion ramp nozzle (SERN) and aircraft-engine integration with an emphasis on the combined cycle engine integration into the conceptual aircraft design. A parametric study using computational fluid dynamics (CFD) have been employed to analyze the sensitivity of the SERN’s performance parameters with changing geometry and operating conditions. The SERN adapted to the different operating conditions and was able to retain its performance throughout the altitude simulated. The expansion ramp shape, angle, exit area, and cowl shape influenced the thrust substantially. The internal nozzle expansion and expansion ramp had a significant effect on the lift and moment performance. An optimized SERN was assembled into a scramjet and was subject to various nozzle inflow conditions, to which combustion flow from twin strut injectors produced the best thrust performance. Side fence studies observed longer and diverging side fences to produce extra thrust compared to small and straight fences. Keywords: scramjet; single expansion ramp nozzle; hypersonic aircraft; combined cycle engines 1.
    [Show full text]
  • Aeronautical Engineering
    NASA/S P--1999-7037/S U P PL407 September 1999 AERONAUTICAL ENGINEERING A CONTINUING BIBLIOGRAPHY WITH INDEXES National Aeronautics and Space Administration Langley Research Center Scientific and Technical Information Program Office The NASA STI Program Office... in Profile Since its founding, NASA has been dedicated CONFERENCE PUBLICATION. Collected to 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 meetings sponsored or cosponsored by NASA. part in helping NASA maintain this important role. SPECIAL PUBLICATION. Scientific, technical, or historical information from The NASA STI Program Office is operated by NASA programs, projects, and missions, Langley Research Center, the lead center for often concerned with subjects having NASA's scientific and technical information. substantial public interest. The NASA STI Program Office provides access to the NASA STI Database, the largest collection TECHNICAL TRANSLATION. of aeronautical and space science STI in the English-language translations of foreign world. The Program Office is also NASA's scientific and technical material pertinent to institutional mechanism for disseminating the NASA's mission. results of its research and development activities. These results are published by NASA in the Specialized services that complement the STI NASA STI Report Series, which includes the Program Office's diverse offerings include following report types: creating custom thesauri, building customized databases, organizing and publishing research TECHNICAL PUBLICATION. Reports of results.., even providing videos. completed research or a major significant phase of research that present the results of For more information about the NASA STI NASA programs and include extensive data or Program Office, see the following: theoretical analysis.
    [Show full text]
  • Sunday Monday
    Sunday Sunday, 29 July 2012 1-RECPT-1 Sunday Opening Reception Exhibit Hall 1830 - 2000 hrs Monday Monday, 30 July 2012 2-JPC-1/IECEC-1 JPC/IECEC Opening Monday Keynote Centennial Ballroom I 0800 - 0900 hrs Overview of NASA major program thrusts and Technology Development Opportunities Robert Lightfoot Associate Administrator NASA Monday, 30 July 2012 3-ABPSI-1/GTE-1 Turboelectric Distributed Propulsion I Hanover C Chaired by: H. KIM, NASA Glenn Research Center and A. GIBSON, Empirical Systems Aerospace LLC 1000 hrs 1030 hrs 1100 hrs 1130 hrs AIAA-2012-3700 AIAA-2012-3701 AIAA-2012-3702 Oral Presentation (Invited) Turboelectric Distributed Sensitivity of Mission Fuel Burn Hybrid Axial and Cross-Flow Evaluation of the Propulsion Propulsion System Modelling to Turboelectric Distributed Fan Propulsion for Transonic Integration Aerodynamics on a for Hybrid-Wing-Body Aircraft Propulsion Design Assumptions Blended Wing Body Aircraft Hybrid Wing Body Concept C. Liu, Self, Cranfield, United on the N3-X Hybrid Wing Body J. Kummer, J. Allred, Propulsive J. Chu, NASA Langley Research Kingdom Aircraft Wing, LLC, Elbridge, NY; J. Felder, Center, Hampton, VA J. Felder, G. Brown, NASA Glenn NASA Glenn Research Center, Research Center, Cleveland, OH; J. Cleveland, OH Chu, NASA Langley Research Center, Hampton, VA; M. Tong, NASA Glenn Research Center, Cleveland, OH Monday, 30 July 2012 4-HSABP/HYP-1/PC-1 Constant Volume Combustion Engines Regency VII Chaired by: V. TANGIRALA, General Electric Company and D. DAUSEN, Naval Postgraduate School 1000 hrs 1030 hrs 1100 hrs 1130 hrs AIAA-2012-3703 AIAA-2012-3704 AIAA-2012-3705 AIAA-2012-3706 Development of a Wave Disk Thermodynamics of the Wave Experimental Optimization of Experimental Study of Shock Engine Experimental Facility Disk Engine Static Valveless Self-Aspiration Transfer in a Multiple Pulse N.
    [Show full text]
  • Flight Data Analysis of the HYSHOT Flight #2 Scramjet
    13th AIAA/CIRA International Space Planes and Hypersonic Systems and Technologies Conference Flight Data Analysis of HyShot 2 Neal E. Hass*, Michael K. Smart+ NASA Langley Research Center, Hampton, Virginia Allan Paull! University of Queensland, Brisbane, Australia Abstract The development of scramjet propulsion for alternative launch and payload delivery capabilities has comprised largely of ground experiments for the last 40 years. With the goal of validating the use of short duration ground test facilities, the University of Queensland, supported by a large international contingency, devised a ballistic re-entry vehicle experiment called HyShot to achieve supersonic combustion in flight above Mach 7.5. It consisted of a double wedge intake and two back-to-back constant area combustors; one supplied with hydrogen fuel at an equivalence ratio of 0.33 and the other un-fueled. Following a first launch failure on October 30th 2001, the University of Queensland conducted a successful second launch on July 30th, 2002. Post-flight data analysis of the second launch confirmed the presence of supersonic combustion during the approximately 3 second test window at altitudes between 35 and 29 km. Reasonable correlation between flight and some pre-flight shock tunnel tests was observed. Nomenclature f aerodynamic factor h altitude M Mach number P pressure q dynamic pressure s seconds T temperature V velocity w mass flow X axial coordinate Y lateral coordinate Z normal coordinate α angle-of-attack ηc combustion efficiency γ ratio of specific heats φ equivelance ratio ω angular velocity about longitudinal body axis ζ angular velocity of longitudinal axis about velocity vector * Aerospace Engineer, Hypersonic Airbreathing Propulsion Branch + Research Scientist, Hypersonic Airbreathing Propulsion Branch ! Professor, Mechanical Engineering Department 1 Subscripts: 0 freestream c combustor entrance w wedge t2 Pitot Introduction The theoretical performance advantage of scramjets over rockets in hypersonic flight has been well known since the 1950’s.
    [Show full text]
  • Ciam/Nasa Mach 6.5 Scramjet Flight and Ground Test
    AIAA-99-4848 CIAM/NASA MACH 6.5 SCRAMJET FLIGHT AND GROUND TEST R. T. Voland* and A. H. Auslender** NASA Langley Research Center Hampton, VA M. K. Smart Lockheed Martin Engineering Sciences Hampton, VA A.S. Roudakovà, V.L. Semenov¤, and V. Kopchenov¤ Central Institute of Aviation Motors Moscow, Russia ABSTRACT NOMENCLATURE The Russian Central Institute of Aviation Motors (CIAM) performed a flight test of a CIAM-designed, C-16V/K Ð CIAM scramjet engine ground test facility, hydrogen-cooled/fueled dual-mode scramjet engine located in Tureavo, Russia (Figure 10) over a Mach number range of approximately 3.5 to 6.4 CIAM Ð Central Institute of Aviation Motors, Moscow, on February 12, 1998, at the Sary Shagan test range in Russia Kazakhstan. This rocket-boosted, captive-carry test of CO2 Ð Carbon dioxide the axisymmetric engine reached the highest Mach Cp Ð Pressure Coefficient number of any scramjet engine flight test to date. The H Ð Altitude (km, m, or ft) flight test and the accompanying ground test program, H2 Ð Hydrogen conducted in a CIAM test facility near Moscow, were H2O Ð Water performed under a NASA contract administered by the He Ð Helium Dryden Flight Research Center with technical HFL Ð Hypersonic Flying Laboratory assistance from the Langley Research Center. Analysis HRE Ð Hypersonic Research Engine of the flight and ground data by both CIAM and NASA HRE AIM Ð Hypersonic Research Engine, resulted in the following preliminary conclusions. An Aerothermodynamic Integration Model unexpected control sensor reading caused non-optimal Hyper-X Ð NASA airframe integrated scramjet- fueling of the engine, and flowpath modifications added powered vehicle flight test program to the engine inlet during manufacture caused markedly ht Ð Total enthalpy (MJ/kg or Btu/lbm) reduced inlet performance.
    [Show full text]
  • Chap21 Rockets Fundamentals.Pdf
    Chapter 21.indd 449 1/9/08 12:12:24 PM There is an explanation for everything that a rocket does. The explanation is most always based on the laws of physics and the nature of rocket propellants. Experimentation is required to find out whether a new rocket will or will not work. Even today, with all the knowledge and expertise that exists in the field of rocketry, experimentation occasionally shows that certain ideas are not practical. In this chapter, we will look back in time to the early developers and users of rocketry. We will review some of the physical laws that apply to rocketry, discuss selected chemicals and their combinations, and identify the rocket systems and their components. We also will look at the basics of rocket propellant efficiency. bjectives Explain why a rocket engine is called a reaction engine. Identify the country that first used the rocket as a weapon. Compare the rocketry advancements made by Eichstadt, Congreve and Hale. Name the scientist who solved theoretically the means by which a rocket could escape the earth’s gravitational field. Describe the primary innovation in rocketry developed by Dr. Goddard and Dr. Oberth. Explain the difference between gravitation and gravity. Describe the contributions of Galileo and Newton. Explain Newton’s law of universal gravitation. State Newton’s three laws of motion. Define force, velocity, acceleration and momentum. Apply Newton’s three laws of motion to rocketry. Identify two ways to increase the thrust of a rocket. State the function of the combustion chamber, the throat, and nozzle in a rocket engine.
    [Show full text]
  • Turbojet Runs Precursor to Hypersonic Engine Heat Exchanger Tests
    Turbojet Runs Precursor to Hypersonic Engine Heat Exchanger Tests Aviation Week & Space Technology Guy Norris Tue, 2018-05-15 04:00 Advanced propulsion developer Reaction Engines is nearing its first step toward validating its novel air-breathing hybrid rocket design at hypersonic conditions by firing up a vintage General Electric J79 turbojet to act as a heat source for testing, expected later this month. The ex-military engine, formerly used in a McDonnell Douglas F-4, is a central element of Reaction’s specially developed high-temperature airflow test site, which will soon be commissioned at Front Range Airport, near Watkins, Colorado. The J79 will provide heated gas flow in excess of 1,000C (1,800F) which, together with conditioned ambient air, will be mixed to replicate inlet conditions representative of flight speeds up to and including Mach 5. The flow will verify the operability and performance of the pre-cooler heat exchanger (HTX), which is at the core of Reaction’s Sabre (synergistic air-breathing rocket engine). It is also key to extracting oxygen from the atmosphere to enable acceleration to hypersonic speed from a standing start. The HTX will chill airflow to minus 150C in less than 1/20th of a second, and pass it through a turbo- compressor and into the rocket combustion chamber where it will be burned with sub-cooled liquid hydrogen (LH) fuel. Beyond Mach 5, and at an altitude approaching 100,000 ft. the inlet will be closed and the engine will continue to operate as a closed-cycle rocket engine fueled by onboard liquid oxygen and LH.
    [Show full text]
  • Investigation of Various Novel Air-Breathing Propulsion Systems
    Investigation of Various Novel Air-Breathing Propulsion Systems A thesis submitted to the Graduate School of the University of Cincinnati in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE in the Department of Aerospace Engineering & Engineering Mechanics of the College of Engineering and Applied Science 2016 by Jarred M. Wilhite B.S. Aerospace Engineering & Engineering Mechanics University of Cincinnati 2016 Committee Chair: Dr. Ephraim Gutmark Committee Members: Dr. Paul Orkwis & Dr. Mark Turner Abstract The current research investigates the operation and performance of various air-breathing propulsion systems, which are capable of utilizing different types of fuel. This study first focuses on a modular RDE configuration, which was mainly studied to determine which conditions yield stable, continuous rotating detonation for an ethylene-air mixture. The performance of this RDE was analyzed by studying various parameters such as mass flow rate, equivalence ratios, wave speed and cell size. For relatively low mass flow rates near stoichiometric conditions, a rotating detonation wave is observed for an ethylene-RDE, but at speeds less than an ideal detonation wave. The current research also involves investigating the newly designed, Twin Oxidizer Injection Capable (TOXIC) RDE. Mixtures of hydrogen and air were utilized for this configuration, resulting in sustained rotating detonation for various mass flow rates and equivalence ratios. A thrust stand was also developed to observe and further measure the performance of the TOXIC RDE. Further analysis was conducted to accurately model and simulate the response of thrust stand during operation of the RDE. Also included in this research are findings and analysis of a propulsion system capable of operating on the Inverse Brayton Cycle.
    [Show full text]
  • Study and Optimization of a Cad/Cfd Model for Valveless Pulsejets
    VOL. 13, NO. 21, NOVEMBER 2018 ISSN 1819-6608 ARPN Journal of Engineering and Applied Sciences ©2006-2018 Asian Research Publishing Network (ARPN). All rights reserved. www.arpnjournals.com STUDY AND OPTIMIZATION OF A CAD/CFD MODEL FOR VALVELESS PULSEJETS Luca Piancastelli1, Stefano Cassani2, Eugenio Pezzuti3 and Luca Lipparini1 1Department of Industrial Engineering, Alma Mater Studiorum University of Bologna, Viale Risorgimento, Bologna (BO), Italy 2MultiProjecta, Via Casola Canina, Imola (BO), Italy 3Università di Roma "Tor Vergata", Dip. di Ingegneria dell’Impresa "Mario Lucertini”, Via del Politecnico, Roma, Italy E-Mail: [email protected] ABSTRACT The method introduced in this paper aims to find a feasible method to evaluate the static thrust of a “valveless” pulsejet, starting from a CAD model. CFD (Computational Fluid Dynamic) simulation and golden section were used for this purpose. Even for new pulsejet designs, it is possible to evaluate the pulsating frequency from equations available in literature or with a mono-dimensional pressure wave model. Then the combustion energy should be introduced in the engine. In this CFD model, the heat flow due to the combustion is simulated through the application of a pulsating flow of hot gases through the walls of the combustion chamber. To minimize the error of this added flow, a stoichiometric combustion of pure oxygen is introduced. The temperature value of the hot gases was optimized with the Golden Section Method in order to obtain the same experimental results of the Department of Aerospace Engineering of California Polytechnic State University, San Luis Obispo [2]. In this way, it is possible to evaluate the performance of a new design of different geometry and size.
    [Show full text]