Fluid Dynamics Android Application: an Efficient Semi-Implicit Solver For
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Fluid Mechanics
cen72367_fm.qxd 11/23/04 11:22 AM Page i FLUID MECHANICS FUNDAMENTALS AND APPLICATIONS cen72367_fm.qxd 11/23/04 11:22 AM Page ii McGRAW-HILL SERIES IN MECHANICAL ENGINEERING Alciatore and Histand: Introduction to Mechatronics and Measurement Systems Anderson: Computational Fluid Dynamics: The Basics with Applications Anderson: Fundamentals of Aerodynamics Anderson: Introduction to Flight Anderson: Modern Compressible Flow Barber: Intermediate Mechanics of Materials Beer/Johnston: Vector Mechanics for Engineers Beer/Johnston/DeWolf: Mechanics of Materials Borman and Ragland: Combustion Engineering Budynas: Advanced Strength and Applied Stress Analysis Çengel and Boles: Thermodynamics: An Engineering Approach Çengel and Cimbala: Fluid Mechanics: Fundamentals and Applications Çengel and Turner: Fundamentals of Thermal-Fluid Sciences Çengel: Heat Transfer: A Practical Approach Crespo da Silva: Intermediate Dynamics Dieter: Engineering Design: A Materials & Processing Approach Dieter: Mechanical Metallurgy Doebelin: Measurement Systems: Application & Design Dunn: Measurement & Data Analysis for Engineering & Science EDS, Inc.: I-DEAS Student Guide Hamrock/Jacobson/Schmid: Fundamentals of Machine Elements Henkel and Pense: Structure and Properties of Engineering Material Heywood: Internal Combustion Engine Fundamentals Holman: Experimental Methods for Engineers Holman: Heat Transfer Hsu: MEMS & Microsystems: Manufacture & Design Hutton: Fundamentals of Finite Element Analysis Kays/Crawford/Weigand: Convective Heat and Mass Transfer Kelly: Fundamentals -
Introduction to Gas Dynamics All Lecture Slides
Introduction to Gas Dynamics AERODYNAMICS - II All Lecture Slides N. Venkata Raghavendra Farheen Sana Autumn 2009 Gasdynamics — Lecture Slides 1 Compressible flow 2 Zeroth law of thermodynamics 3 First law of thermodynamics 4 Equation of state — ideal gas 5 Specific heats 6 The “perfect” gas 7 Second law of thermodynamics 8 Adiabatic, reversible process 9 The free energy and free enthalpy 10 Entropy and real gas flows 11 One-dimensional gas dynamics 12 Conservation of mass — continuity equation 13 Conservation of energy — energy equation 14 Reservoir conditions 15 On isentropic flows Gasdynamics — Lecture Slides 16 Euler’s equation 17 Momentum equation 18 A review of equations of conservation 19 Isentropic condition 20 Speed of sound — Mach number 21 Results from the energy equation 22 The area-velocity relationship 23 On the equations of state 24 Bernoulli equation — dynamic pressure 25 Constant area flows 26 Shock relations for perfect gas — Part I 27 Shock relations for perfect gas — Part II 28 Shock relations for perfect gas — Part III 29 The area-velocity relationship — revisited 30 Nozzle flow — converging nozzle Gasdynamics — Lecture Slides 31 Nozzle flow — converging-diverging nozzle 32 Normal shock recovery — diffuser 33 Flow with wall roughness — Fanno flow 34 Flow with heat addition — Rayleigh flow 35 Normal shock, Fanno flow and Rayleigh flow 36 Waves in supersonic flow 37 Multi-dimensional equations of the flow 38 Oblique shocks 39 Relationship between wedge angle and wave angle 40 Small angle approximation 41 Mach lines 42 Weak oblique -
Chapter Sixteen / Plug, Underexpanded and Overexpanded Supersonic Nozzles ------Chapter Sixteen / Plug, Underexpanded and Overexpanded Supersonic Nozzles
UOT Mechanical Department / Aeronautical Branch Gas Dynamics Chapter Sixteen / Plug, Underexpanded and Overexpanded Supersonic Nozzles -------------------------------------------------------------------------------------------------------------------------------------------- Chapter Sixteen / Plug, Underexpanded and Overexpanded Supersonic Nozzles 16.1 Exit Flow for Underexpanded and Overexpanded Supersonic Nozzles The variation in flow patterns inside the nozzle obtained by changing the back pressure, with a constant reservoir pressure, was discussed early. It was shown that, over a certain range of back pressures, the flow was unable to adjust to the prescribed back pressure inside the nozzle, but rather adjusted externally in the form of compression waves or expansion waves. We can now discuss in detail the wave pattern occurring at the exit of an underexpanded or overexpanded nozzle. Consider first, flow at the exit plane of an underexpanded, two-dimensional nozzle (see Figure 16.1). Since the expansion inside the nozzle was insufficient to reach the back pressure, expansion fans form at the nozzle exit plane. As is shown in Figure (16.1), flow at the exit plane is assumed to be uniform and parallel, with . For this case, from symmetry, there can be no flow across the centerline of the jet. Thus the boundary conditions along the centerline are the same as those at a plane wall in nonviscous flow, and the normal velocity component must be equal to zero. The pressure is reduced to the prescribed value of back pressure in region 2 by the expansion fans. However, the flow in region 2 is turned away from the exhaust-jet centerline. To maintain the zero normal-velocity components along the centerline, the flow must be turned back toward the horizontal. -
Normal and Oblique Shocks, Prandtl Meyer Expansion
Aerothermodynamics of high speed flows AERO 0033{1 Lecture 3: Normal and oblique shocks, Prandtl Meyer expansion Thierry Magin, Greg Dimitriadis, and Adrien Crovato [email protected] Aeronautics and Aerospace Department von Karman Institute for Fluid Dynamics Aerospace and Mechanical Engineering Department Faculty of Applied Sciences, University of Li`ege Wednesday 9am { 12:15pm February { May 2021 1 / 36 Outline Normal shock Total and critical quantities Oblique shock Prandtl-Meyer expansion Shock reflection and shock interaction 2 / 36 Outline Normal shock Total and critical quantities Oblique shock Prandtl-Meyer expansion Shock reflection and shock interaction 2 / 36 Normal shock relations (inviscid flow) Rankine-Hugoniot relations for steady normal shock σ = 0; v = u ex ; n = ex ! σn(U2 − U1) = (F · n)2 − (F · n)1 ρ2u2 = ρ1u1 2 2 ρ2u2 + p2 = ρ1u1 + p1 ρ2u2H2 = ρ1u1H1 I Eqs. also valid for non calorically perfect gases I Contact discontinuity when u2 = u1 = 0 ! p2 = p1 I Otherwise, the total enthalpy conservation can be expressed as 2 2 2 u cp p γ p2 u2 γ p1 u1 H2 = H1 with H = h+ ; h = ! + = + 2 R ρ γ − 1 ρ2 2 γ − 1 ρ1 2 I Non-linear algebraic system of 3 eqs. in 3 unknowns ρ2, u2, and p2 with closed solution expressed in function of dimensionless parameters: Machp number M1 = u1=a1 and specific heat ratio γ, with a1 = γRT1 3 / 36 Normal shock relations for calorically perfect gases For M1 > 1 (derivation given further in this section) 2 ρ2 u1 (γ+1)M1 I ρ = u = 2 > 1 1 2 2+(γ−1)M1 p2 = 1 + 2γ (M2 − 1) > 1 I p1 γ+1 1 γ−1 2 2 1+ 2 -
Introduction
CHAPTER 1 Introduction "For some years I have been afflicted with the belief that flight is possible to man." Wilbur Wright, May 13, 1900 1.1 ATMOSPHERIC FLIGHT MECHANICS Atmospheric flight mechanics is a broad heading that encompasses three major disciplines; namely, performance, flight dynamics, and aeroelasticity. In the past each of these subjects was treated independently of the others. However, because of the structural flexibility of modern airplanes, the interplay among the disciplines no longer can be ignored. For example, if the flight loads cause significant structural deformation of the aircraft, one can expect changes in the airplane's aerodynamic and stability characteristics that will influence its performance and dynamic behavior. Airplane performance deals with the determination of performance character- istics such as range, endurance, rate of climb, and takeoff and landing distance as well as flight path optimization. To evaluate these performance characteristics, one normally treats the airplane as a point mass acted on by gravity, lift, drag, and thrust. The accuracy of the performance calculations depends on how accurately the lift, drag, and thrust can be determined. Flight dynamics is concerned with the motion of an airplane due to internally or externally generated disturbances. We particularly are interested in the vehicle's stability and control capabilities. To describe adequately the rigid-body motion of an airplane one needs to consider the complete equations of motion with six degrees of freedom. Again, this will require accurate estimates of the aerodynamic forces and moments acting on the airplane. The final subject included under the heading of atmospheric flight mechanics is aeroelasticity. -
Nozzle Aerodynamics Baseline Design
Preliminary Design Review Supersonic Air-Breathing Redesigned Engine Nozzle Customer: Air Force Research Lab Advisor: Brian Argrow Team Members: Corrina Briggs, Jared Cuteri, Tucker Emmett, Alexander Muller, Jack Oblack, Andrew Quinn, Andrew Sanchez, Grant Vincent, Nathaniel Voth Project Description Model, manufacture, and verify an integrated nozzle capable of accelerating subsonic exhaust to supersonic exhaust produced from a P90-RXi JetCat engine for increased thrust and efficiency from its stock configuration. Stock Nozzle Vs. Supersonic Nozzle Inlet Compressor Combustor Turbine Project Baseline Nozzle Nozzle Test Nozzle Project Description Design Aerodynamics Bed Integration Summary Objectives/Requirements •FR 1: The Nozzle Shall accelerate the flow from subsonic to supersonic conditions. •FR 2: The Nozzle shall not decrease the Thrust-to-Weight Ratio. •FR 3: The Nozzle shall be designed and manufactured such that it will integrate with the JetCat Engine. •DR 3.1: The Nozzle shall be manufactured using additive manufacturing. •DR 3.4: Successful integration of the nozzle shall be reversible such that the engine is operable in its stock configuration after the new nozzle has been attached, tested, and detached. •FR4: The Nozzle shall be able to withstand engine operation for at least 30 seconds. Project Baseline Nozzle Nozzle Test Nozzle Project Description Design Aerodynamics Bed Integration Summary Concept of Operations JetCat P90-SE Subsonic Supersonic Engine Flow Flow 1. Remove Stock Nozzle 2. Additive Manufactured 3-D Nozzle -
Gas Dynamics and Jet Propulsion
A Course Material on GAS DYNAMICS AND JET PROPULSION By Mr. C.RAVINDIRAN. ASSISTANT PROFESSOR DEPARTMENT OF MECHANICAL ENGINEERING SASURIE COLLEGE OF ENGINEERING VIJAYAMANGALAM – 638 056 QUALITY CERTIFICATE This is to certify that the e-course material Subject Code : ME2351 Subject : GAS DYNAMICS AND JET PROPULSION Class : III Year MECHANICAL ENGINEERING being prepared by me and it meets the knowledge requirement of the university curriculum. Signature of the Author Name : C. Ravindiran Designation: Assistant Professor This is to certify that the course material being prepared by Mr. C. Ravindiran is of adequate quality. He has referred more than five books amount them minimum one is from abroad author. Signature of HD Name: Mr. E.R.Sivakumar SEAL CONTENTS S.NO TOPIC PAGE NO UNIT-1 BASIC CONCEPTS AND ISENTROPIC 1.1 Concept of Gas Dynamics 1 1.1.1 Significance with Applications 1 1.2 Compressible Flows 1 1.2.1 Compressible vs. Incompressible Flow 2 1.2.2 Compressibility 2 1.2.3. Compressibility and Incompressibility 3 1.3 Steady Flow Energy Equation 5 1.4 Momentum Principle for a Control Volume 5 1.5 Stagnation Enthalpy 5 1.6 Stagnation Temperature 7 1.7 Stagnation Pressure 8 1.8 Stagnation velocity of sound 9 1.9 Various regions of flow 9 1.10 Flow Regime Classification 11 1.11 Reference Velocities 12 1.11.1 Maximum velocity of fluid 12 1.11.2 Critical velocity of sound 13 1.12 Mach number 16 1.13 Mach Cone 16 1.14 Reference Mach number 16 1.15 Crocco number 19 1.16 Isothermal Flow 19 1.17 Law of conservation of momentum 19 1.17.1 Assumptions -
Experimental Investigation Into Shock Waves Formation and Development Process in Transonic Ow
Scientia Iranica B (2017) 24(5), 2457{2465 Sharif University of Technology Scientia Iranica Transactions B: Mechanical Engineering www.scientiairanica.com Experimental investigation into shock waves formation and development process in transonic ow M. Farahani and A. Jaberi Department of Aerospace Engineering, Sharif University of Technology, Tehran, Iran. Received 11 May 2016; received in revised form 12 August 2016; accepted 8 October 2016 KEYWORDS Abstract. An extensive experimental investigation was performed to explore the shock waves formation and development process in transonic ow. Shadowgraph visualization Shock waves technique was employed to provide visual description of the ow eld features. Based on formation process; the visualization, the formation process was categorized into two intrinsically di erent Transonic ow; phases, namely, subsonic and supersonic. The characteristics of subsonic phase are well Shadowgraphy known; however, those of the supersonic one are far less studied. The supersonic phase visualization; itself is made up of two consecutive phases, namely, approaching and sweeping. The e ects Splitting of shock of each phase on the ow eld characteristics and on shaping the supersonic regime were waves; studied in details. In order to generalize the results, three di erent models were tested. A Zeh shock. special terminology is suggested by authors to ease the process description and pave a way for future studies. Above all, as the transition from transonic regime to the supersonic one is a vague concept in terms of physical reasoning, a new explanation is proposed that can be used as a criterion for distinguishing between transonic and supersonic regimes. © 2017 Sharif University of Technology. -
Bifurcating Mach Shock Reflections with Application to Detonation Structure
BIFURCATING MACH SHOCK REFLECTIONS WITH APPLICATION TO DETONATION STRUCTURE Philip Mach A thesis submitted to the Faculty of Graduate and Postdoctoral Studies in partial fulfillment of the requirements for the degree of MASTER OF APPLIED SCIENCE in Mechanical Engineering Ottawa-Carleton Institute for Mechanical and Aerospace Engineering University of Ottawa Ottawa, Canada August 2011 © Philip Mach, Ottawa, Canada, 2011 ii Abstract Numerical simulations of Mach shock reflections have shown that the Mach stem can bifurcate as a result of the slip line jetting forward. Numerical simulations were conducted in this study which determined that these bifurcations occur when the Mach number is high, the ramp angle is high, and specific heat ratio is low. It was clarified that the bifurcation is a result of a sufficiently large velocity difference across the slip line which drives the jet. This bifurcation phenomenon has also been observed after triple point collisions in detonation simulations. A triple point reflection was modelled as an inert shock reflecting off a wedge, and the accuracy of the model at early times after reflection indicates that bifurcations in detonations are a result of the shock reflection process. Further investigations revealed that bifurcations likely contribute to the irregular structure observed in certain detonations. iii Acknowledgements Firstly, I would like to thank my supervisor, Matei Radulescu, who guided me through the complicated task of learning about the detonation phenomenon, provided me with most of the scripts for the numerical simulations, answered my questions (no matter how dumb they were), and was patient with my progress. I would have obviously been lost without him. -
Investigation of Shock Wave Interactions Involving Stationary
Investigation of Shock Wave Interactions involving Stationary and Moving Wedges Pradeep Kumar Seshadri and Ashoke De* Department of Aerospace Engineering, Indian Institute of Technology Kanpur, Kanpur, 208016, India. *Corresponding Author: [email protected] Abstract The present study investigates the shock wave interactions involving stationary and moving wedges using a sharp interface immersed boundary method combined with a fifth-order weighted essentially non- oscillatory (WENO) scheme. Inspired by Schardin’s problem, which involves moving shock interaction with a finite triangular wedge, we study influences of incident shock Mach number and corner angle on the resulting flow physics in both stationary and moving conditions. The present study involves three incident shock Mach numbers (1.3, 1.9, 2.5) and three corner angles (60°, 90°, 120°), while its impact on the vorticity production is investigated using vorticity transport equation, circulation, and rate of circulation production. Further, the results yield that the generation of the vorticity due to the viscous effects are quite dominant compared to baroclinic or compressibility effects. The moving cases presented involve shock driven wedge problem. The fluid and wedge structure dynamics are coupled using the Newtonian equation. These shock driven wedge cases show that wedge acceleration due to the shock results in a change in reflected wave configuration from Single Mach Reflection (SMR) to Double Mach Reflection (DMR). The intermediary state between them, the Transition Mach Reflection (TMR), is also observed in the process. The effect of shock Mach number and corner angle on Triple Point (TP) trajectory, as well as on the drag coefficient, is analyzed in this study. -
Numerical Study of Interaction of a Vortical Density Inhomogeneity with Shock and Expansion Waves
22/ NASA/CR- 1998-206918 ICASE Report No. 98-10 Numerical Study of Interaction of a Vortical Density Inhomogeneity with Shock and Expansion Waves A. Povitsky ICASE D. Ofengeim University of Manchester Institute of Science and Technology Institute for Computer Applications in Science and Engineering NASA Langley Research Center Hampton, VA Operated by Universities Space Research Association National Aeronautics and Space Administration Langley Research Center Prepared for Langley Research Center under Contract NAS 1-97046 Hampton, Virginia 23681-2199 February 1998 Available from the following: NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS) 800 Elkridge Landing Road 5285 Port Royal Road Linthicum Heights, MD 21090-2934 Springfield, VA 22161-2171 (301) 621-0390 (703) 487-4650 NUMERICAL STUDY OF INTERACTION OF A VORTICAL DENSITY INHOMOGENEITY WITH SHOCK AND EXPANSION WAVES A. POVITSKY * AND D. OFENGEIM t Abstract. We studied the interaction of a vortical density in_homogeneity (VDI) with shock and expan- sion waves. We call the VDI the region of concentrated vorticity (vortex) with a density different from that of ambiance. Non-parallel directions of the density gradient normal to the VDI surface and the pressure gradient across a shock wave results in an additional vorticity. The roll-up of the initial round VDI towards a non-symmetrical shape is studied numerically. Numerical modeling of this interaction is performed by a 2-D Euler code. The use of an adaptive unstructured numerical grid makes it possible to obtain high accuracy and capture regions of induced vorticity with a moderate overall number of mesh points. For the validation of the code, the computational results are compared with available experimental results and good agreement is obtained. -
Copyrighted Material
BBMIndx.inddMIndx.indd PPageage I-1I-1 111/10/181/10/18 6:536:53 PMPM f-389f-389 //208/WB02304_EVALCS/9781119469582/bmmatter/text_s208/WB02304_EVALCS/9781119469582/bmmatter/text_s Index Absolute pressure, 13, 41 Angular momentum flow, 174, video V5.12 restrictions on use of, 106–110, 222, Absolute temperature Annulus, flow in, 257–258 video V3.12 in British Gravitational (BG) System, 7 Apparent viscosity, 16 unsteady flow, 107–108 in English Engineering (EE) System, 8 Aqueducts, 454 use of, video V3.2 in ideal gas law, 12 Archimedes, 27, 28, 62 Best efficiency points, 564 in International System (SI), 7 Archimedes’ principle, 61–64, video V2.7 Best hydraulic cross section, 460–461 Absolute velocity Area BG units. see British Gravitational (BG) System in centrifugal pumps, 558–561 centroid, 51–53 of units defined, 154, 551 compressible flow with change in, 507–522 Bingham plastic, 16–17 in rotating sprinkler example, 176–177 isentropic flow of ideal gas with change in, Birds, wings of, 429 in turbo machine example, 551–554 510–516 Blades vs. relative velocity, 141–142 and Mach number, 511–513 curved and radia, 561, 562 Absolute viscosity moment of inertia, 52 radial, 561 of air, 620–621 Aspect ratio rotor, 575, 591 of common gases, 617 air foil, 435 stator, 575, 591 of common liquids, 617 plate, 271 turbo machine, 549 defined, 15 Assemblies, 603 Blade velocity, 551–554, 558–559 of U.S. Standard Atmosphere, 622–623 Atmosphere, standard. see Standard Blasius, H., 29, 285, 389 of water, 619 atmosphere Blasius formula, 285, 339 Acceleration