Fundamentals of Pulse Detonation Engine (PDE) and Related Propulsion Technology

Transcription

1 Fundamentals of Pulse Detonation Engine (PDE) and Related Propulsion Technology Dora E. Musielak, Ph.D. Aerospace Engineering Consulting Arlington, TX All rights reserved. No part of this publication may be reproduced, distributed, or transmitted, unless for course participation, in any form or by any means, or stored in a database or retrieval system, without the prior written permission of the Author. Contact D. E. Musielak, dmusielak@uta.edu

2 Pure PDE Cycle Pulse Detonation Engine (PDE) : type of propulsion system that utilizes detonation waves to combust fuel and oxidizer mixture. Engine is pulsed because mixture must be renewed from combustion chamber between each detonation wave initiated. 1: Fuel-Oxidizer Injected and Mixed 2: Detonation Initiated by ignition source 3: Detonation wave moves through gas mixture 4: High pressure gas fills detonation chamber 5: Detonation wave exits chamber and air is drawn in by reduced pressure FILL DETONATE EXHAUST Repeat Thrust is directly proportional to detonation frequency

3 Why PDEs? Advantages Increased Thermodynamic Efficiency Higher Isp Reduced SFC Design Simplicity Increased Thrust-to-Weight Ratio Increased Thrust-to-Volume Ratio Lower Cost Mach Range 0 4 Easy Vehicle Integration Applications Cruise Missiles Supersonic Aircraft Hypersonic Missiles Hybrid Turbine-PDE UAV UCAV SSTO Launch Vehicles Precision Guided Munitions Drones Increased cycle efficiency results from quasi-constant volume process PDEs potential for easier scaling extrapolates to substantial reductions in development time, when compared to conventional turbine engines. 3

4 Why PDEs? Is it possible to augment gas turbine performance with PDEs to extend supersonic flight regime? 4

5 Why PDEs? Is a PDE/SCRAM/PDRE a viable propulsion system for Spaceplanes? PDRE = Pulse Detonation Rocket Engine SCRAM = Scramjet supersonic combustion ramjet engine for M > 5 5

6 Preface Revolutionary propulsion is required to achieve high-speed cruise capability within atmosphere, and for low cost reliable Earth-to-orbit vehicles. Pulse detonation engines (PDEs) have potential performance advantages over air breathing and rocket propulsion, bypassing limitations of existing concepts. Proposed applications for detonation combustion include cruise missiles, UAV,... supersonic aircraft, and SSTO launchers. This course highlights fundamentals of pulse detonation engines and other related propulsion concepts, addressing performance characteristics, enabling technologies, and current R&D initiatives to develop new propulsion systems. 6

7 Nomenclature and Terminology A list of common terms and basic definitions is provided in a separate handout to facilitate communicating the concepts introduced in the course. In 2002, Kaemming, Lidstone, and Sammann proposed a component nomenclature, station (spatial) designation, process and event (temporal) designation and terminology for the unique PDE scheduling characteristics. Ref. AIAA Nomenclature proposal is based on several years of PDE analysis and testing by Boeing and Pratt & Whitney and is based on accepted practices, such as SAE Standard AS7551. To date, no standard has been formally issued for PDEs, and so we will follow the recommendations in AIAA paper

8 Nomenclature and Terminology See Appendix 1 8

9 Introduction to PDEs Propulsion Comparison A Vision for the Future Limits of Turbo-Engines Ideal Cycles Combustion Modes Pure Pulse Detonation Engine Detonation for Propulsion Modeling a Single Cycle PDE Thermodynamic Cycle 9

10 Seeking Revolutionary Propulsion Ideas Propulsion Comparison I sp F g m o f SFC m F f Need improved SFC performance 10

11 Highest Supersonic Speed: M = 3.2 Turbofan engines in high performance aircraft such as F-15 Eagle fighter can achieve Mach 2.5 F-16 Fighting Falcon jet fighter and F-22 Raptor are limited to Mach 2. SR-71 11

12 A Vision for the Future Air Breathing Propulsion Requirements Reduced SFC Higher Thermodynamic Efficiency Higher Isp Design Simplicity Increased Thrust-to-Weight Ratio Increased Thrust-to-Volume Ratio Lower Cost Mach range 0 10 Manufacturing Simplicity Easy Vehicle Integration Develop air-breathing engine capable of propelling aircraft beyond Mach

13 Turbine Engine Limits At Mach > 3, compressed air reaches such extreme temperatures that compressor stage fan blades begin to fail. Compressor exit temperature limits pressure ratio. Turbine inlet temperature limits thrust. Inlet Efficiency Static pressure balance P&W F100 AB Turbofan Rotational speed Compressor exit temperature limits pressure ratio F ma o 2 Tt ( 4 1) 1 T o 13

14 Turbofan with Afterburner Efficient with continuous afterburner at ~ Mach 3. Afterburner provides temporary increase in thrust, for supersonic flight and take off Slower bypass airflow produces thrust more efficiently than high-speed air from core, reducing specific fuel consumption. 14

15 P&W F-100- Performance Maximum thrust: 17,800 lbf (79.1 kn) military thrust 29,160 lbf (129.6 kn) with afterburner Overall pressure ratio: 32:1 F-16 Specific fuel consumption: Military thrust: 0.76 lb/(lbf h) (77.5 kg/(kn h)) Full afterburner: 1.94 lb/(lbf h) (197.8 kg/(kn h)) Thrust-to-weight ratio: 7.8:1 (76.0 N/kg) 15

16 Ideal Thermodynamic Cycle Brayton cycle: heat addition at constant pressure Q in netw mc out p W t ( T ) 4 T3 W c mc p [ T4 T5 ( T3 T2 )] Q in th netw Q in out = rate of thermal energy released netw out = net power out of engine th = thermal efficiency of engine B ( T T ) ( T T ) ( T T ) ( T ( T 4 T ) T ) 3 T T B 1 ( 1) / ( p3 / p2) 16

17 Burner Exit Temperature T4 Increasing T4 enlarges useful work output (isobars diverge ) However, distance between stations 3 and 4 increases also more heat has to be added and thus more fuel is needed. Thermal efficiency is only dependant on compressor pressure ratio P3/P2 and does not change with T4 17

18 Higher T4, Lower Efficiency However, isentropic exponent is not constant but decreases when temperature increases thermodynamic efficiency decreases with T4! 18

19 Heat Addition and Pressure 4 Can we improve thermodynamic cycle efficiency with a pressure-gain process? 19

20 Ideal Thermodynamic Cycle 3* Constant volume heat addition Q in 3 4 Humphrey cycle: heat addition at constant volume 2 5 H T 1 T T 4 1 T3 T4 1 T Thermal efficiency improves by more than 15% and as much as 10 to 40% improvement in Isp (Ref. Bussing (1996); Heiser & Pratt (2002); Povinelli (2002)) 20

21 Pulse Detonation Engine (PDE) Ideal PDE thermodynamic efficiency higher than turboengine because a detonation wave rapidly compresses mixture and adds heat at ~ constant volume. 21

22 Propulsion Performance 22

23 Combustion Modes F Deflagration Subsonic Combustion Detonation Supersonic Combustion Unsteady Pulsed or Intermittent Steady or Continuous Unsteady Pulsed or Intermittent Steady or Continuous Combustion Pulse Jets Turbojets Ramjets Scramjets Deflagration subsonic spread of combustion by thermal conductivity PDEs PDREs Scramjets? RDE Detonation supersonic spread of combustion by shock compression. 23

24 u 1 u 2 Combustion Modes: Detonation and Deflagration Deflagrations are subsonic combustion waves: M 1 < 1 Typical deflagrations propagate at speeds on the order of m/s Across a deflagration, the pressure decreases while the volume increases, P 2 < P 1 and V 2 > V 1 Detonations are supersonic waves: M 1 >1 Typical detonation waves propagate at a velocity on the order of 2000 m/s (4 < M 1 < 8) Pressure increase across a detonation, while the volume decreases: P 2 > P 1, V 2 < V 1 Detonations in HC fuel: P 2 /P 1 ~ 20 P 1 P 2 24

25 Deflagration and Detonation Combustion or burning is sequence of exothermic chemical reactions between fuel and an oxidant accompanied by production of heat and conversion of chemical species. Flame propagates from right Reactants u 1 P 1, T 1, 1, M 1 Products u 2 P 2, T 2, 2, M 2 25

26 Detonation vs Deflagration: Qualitative Differences Qualitative differences between upstream and downstream properties across detonation wave are similar to property differences across normal shock Main differences: Normal shock wave: downstream velocity always subsonic Detonation wave: downstream velocity always local speed of sound Note that detonation waves can fall into strong and weak classes Strong detonation: subsonic burned gas velocity Weak detonation: supersonic burned gas velocity 26

27 Wave Properties Normal shock property ratios are qualitatively similar to those of detonations and of same magnitude Except that for detonation downstream velocity is sonic Mach number increases across flame for deflagrations Mach number is very small and thus is not a very useful parameter to characterize a deflagration Velocity increases substantially and density drops substantially across a deflagration Effects are opposite in direction as compared with detonations or shock waves Pressure is essentially constant across a deflagration (actually it decreases slightly), while detonation has high pressure downstream of propagating wave Characteristic shared by shock, detonation, and deflagration is large temperature increase across wave 27

28 Detonation wave (DW) propagation to create thrust Detonation for Propulsion Oblique Detonation Wave Engine (ODWE) Combustible gas mixture velocity equals or exceeds detonation Chapman- Jouguet (CJ) velocity. Detonation waves (DWs) or oblique detonation waves (ODWs) are positioned to combust injected combustible mixture. Pulse Detonation Engine (PDE) Cyclically detonates fuel and atmospheric air mixtures to generate thrust. A shock wave compresses gas and this is followed by rapid release of heat and a sudden rise in pressure. PDE generates thrust intermittently, and it produces a significant pressure rise in combustor. Detonation-generated pressure rise represents primary benefits of a PDE in that it may reduce engine compression requirements. Continuous Detonation Engine (CDE) Combustible gas mixture is injected along axial direction, and DWs propagate in azimuthal direction. Two directions are independent, DWs can continuously propagate with range of combustible gas injection velocities and do not require multi-time ignition. 28

29 Pure PDE Cycle 1: Fuel-Oxidizer Injected and Mixed 2: High pressure detonation Initiated 3: Detonation wave moves through gas mixture at supersonic speed 4: High pressure gas fills detonation chamber 5: Detonation wave exits chamber and air is drawn in by reduced pressure FILL DETONATE Repeat EXHAUST Thrust is directly proportional to detonation frequency

30 Detonation Initiation A detonation may form via direct initiation or via deflagration-todetonation transition (DDT). Direct initiation is dependent upon an ignition source driving a blast wave of sufficient strength such that igniter is directly responsible for initiating detonation. It requires extremely large energy. DDT begins with a deflagration initiated by relatively weak energy source which accelerates through interactions with its surroundings into a coupled shock wave-reaction zone structure characteristic of a detonation. After spark creates a deflagration, transition process can take several meters or longer and a large amount of time. Key to detonation initiation schemes for PDEs is to shorten distance and time required for deflagration-to-detonation transition (DDT). 30

31 PDE Requirements Ignition and mixing must occur quickly to minimize cycle time and maximize thrust. DDT must occur quickly and in a short distance. Shortening DDT time decreases the detonate part of the cycle, allowing a frequency increase that is accompanied by a thrust increase. Shortening DDT distance decreases necessary thrust tube length, resulting in weight savings, a great advantage for propulsion

32 PDE Cycle (Basic Cycle Process) 1. Initially, chamber at ambient conditions 2. Propellant injected from closed end * Sidewall injection also works and may improve mixing 3. Ignition from closed end 4. Wave propagation and transition in chamber 5. Wave exits chamber 6. Exhaust and purge 32

33 Rankine-Hugoniot Combustion Map Conservation equations for mass, momentum, and energy for combustion waves in steady, inviscid, and constant-area flow. P 1v1 2v v1 P2 2v2 M 1 M 2 P1, v1, P 1 2, v2, 2 Combustion wave h v1 h2 2 v2 Hugoniot is locus of possible solutions for state 2 from a given state 1 and a given energy release Rayleigh line relates states 1 and 2. Solution state is at intersection of Hugoniot and Rayleigh line. 33

34 Chapman-Jouguet C-J Condition Solution to conservation equations is determined considering: For deflagrations, wave structure, and turbulent and diffusive processes determine propagation speed. For detonations, gas dynamic considerations are sufficient to determine solution. Chapman (1899) and Jouguet (1905) proposed that detonations travel at one particular velocity, which is minimum velocity for all solutions on detonation branch. At solution point (Chapman-Jouguet detonation point), Hugoniot, Rayleigh line, and isentrope are tangent. Flow behind a C-J detonation is sonic relative to wave: M 2 =1. C-J points divide Hugoniot into 4 regions: Weak deflagrations (subsonic to subsonic) Strong deflagrations (subsonic to supersonic) Weak detonations (supersonic to supersonic) Strong detonations (supersonic to subsonic) 34

35 C-J Velocity Chapman-Jouguet (C-J) condition: for an infinitesimal thin detonation, detonation wave proceeds at a velocity at which reacting gases just reach sonic velocity (in frame of lead shock) as reaction ceases. Assumes chemical reaction takes place at moment when shock compresses material Chapman-Jouguet velocity: velocity of an ideal detonation as determined by C-J condition: burned gas at end of reaction zone travels at sound speed relative to detonation wave front. C-J velocities can be computed numerically by solving for thermodynamic equilibrium and satisfying mass, momentum, and energy conservation for a steadily-propagating wave terminating in a sonic point. C-J velocities in typical fuel-air mixtures between 1400 and 1800 m/s. Speed of sound: 331 m/s in air. 35

36 PDE Thermodynamic Cycle (Heiser & Pratt, 2002 Process 3 4 models normal detonation wave in a PDE (ZND wave model) Entropy generated in detonation wave heat addition process is sum of that generated in process from 3 to 3a (adiabatic normal shock wave) and that generated in process from 3a to 4 (constant-area heat addition process) that follows. Thermal efficiency of ideal Humphrey cycle is close to, but always somewhat less than, that of ideal PDE cycle H&P 36

37 Summary of Chapter 1 Revolutionary propulsion is required to achieve high-speed cruise capability within atmosphere, and for low cost reliable earth-to-orbit vehicles. Pulse detonation engines (PDEs) have potential performance advantages over air breathing and rocket propulsion, bypassing limitations of existing concepts. Propulsion architectures that use pulsed and continuous detonation combustion offer more efficient thermodynamic properties, and thus are expected to exhibit a higher level of performance than more conventional propulsion that rely simply on deflagration combustion process. Chapter 2 will provide an overview of detonation-based propulsion, including hybrid turbine-pde and Continuous Detonation Wave Engine (CDWE) concepts. 37

38 J58 R-R/Snecma Olympus 593 P&W F100 GE F110 P&W F119 GE F414 P&W F Next Generation Supersonic Air Breathing Engine 38

39 Detonation Detonation is a shock wave sustained by energy released by combustion Combustion process, in turn, is initiated by shock wave compression and resulting high temperatures Detonations involve interaction between fluid mechanic processes (shock waves) and thermochemical processes (combustion) 39