An Introduction to Aerodynamics
Total Page:16
File Type:pdf, Size:1020Kb
Load more
Recommended publications
-
Notes on Earth Atmospheric Entry for Mars Sample Return Missions
NASA/TP–2006-213486 Notes on Earth Atmospheric Entry for Mars Sample Return Missions Thomas Rivell Ames Research Center, Moffett Field, California September 2006 The NASA STI Program Office . in Profile Since its founding, NASA has been dedicated to the • CONFERENCE PUBLICATION. Collected advancement of aeronautics and space science. The papers from scientific and technical confer- NASA Scientific and Technical Information (STI) ences, symposia, seminars, or other meetings Program Office plays a key part in helping NASA sponsored or cosponsored by NASA. maintain this important role. • SPECIAL PUBLICATION. Scientific, technical, The NASA STI Program Office is operated by or historical information from NASA programs, Langley Research Center, the Lead Center for projects, and missions, often concerned with NASA’s scientific and technical information. The subjects having substantial public interest. NASA STI Program Office provides access to the NASA STI Database, the largest collection of • TECHNICAL TRANSLATION. English- aeronautical and space science STI in the world. language translations of foreign scientific and The Program Office is also NASA’s institutional technical material pertinent to NASA’s mission. mechanism for disseminating the results of its research and development activities. These results Specialized services that complement the STI are published by NASA in the NASA STI Report Program Office’s diverse offerings include creating Series, which includes the following report types: custom thesauri, building customized databases, organizing and publishing research results . even • TECHNICAL PUBLICATION. Reports of providing videos. completed research or a major significant phase of research that present the results of NASA For more information about the NASA STI programs and include extensive data or theoreti- Program Office, see the following: cal analysis. -
Introduction to Aerodynamics < 1.7. Dimensional Analysis > Physical
Introduction to Aerodynamics < 1.7. Dimensional analysis > Physical parameters Q: What physical quantities influence forces and moments? V A l Aerodynamics 2015 fall - 1 - Introduction to Aerodynamics < 1.7. Dimensional analysis > Physical parameters Physical quantities to be considered Parameter Symbol units Lift L' MLT-2 Angle of attack α - -1 Freestream velocity V∞ LT -3 Freestream density ρ∞ ML -1 -1 Freestream viscosity μ∞ ML T -1 Freestream speed of sound a∞ LT Size of body (e.g. chord) c L Generally, resultant aerodynamic force: R = f(ρ∞, V∞, c, μ∞, a∞) (1) Aerodynamics 2015 fall - 2 - Introduction to Aerodynamics < 1.7. Dimensional analysis > The Buckingham PI Theorem The relation with N physical variables f1 ( p1 , p2 , p3 , … , pN ) = 0 can be expressed as f 2( P1 , P2 , ... , PN-K ) = 0 where K is the No. of fundamental dimensions Then P1 =f3 ( p1 , p2 , … , pK , pK+1 ) P2 =f4 ( p1 , p2 , … , pK , pK+2 ) …. PN-K =f ( p1 , p2 , … , pK , pN ) Aerodynamics 2015 fall - 3 - < 1.7. Dimensional analysis > The Buckingham PI Theorem Example) Aerodynamics 2015 fall - 4 - < 1.7. Dimensional analysis > The Buckingham PI Theorem Example) Aerodynamics 2015 fall - 5 - < 1.7. Dimensional analysis > The Buckingham PI Theorem Aerodynamics 2015 fall - 6 - < 1.7. Dimensional analysis > The Buckingham PI Theorem Through similar procedure Aerodynamics 2015 fall - 7 - Introduction to Aerodynamics < 1.7. Dimensional analysis > Dimensionless form Aerodynamics 2015 fall - 8 - Introduction to Aerodynamics < 1.8. Flow similarity > Dynamic similarity Two different flows are dynamically similar if • Streamline patterns are similar • Velocity, pressure, temperature distributions are same • Force coefficients are same Criteria • Geometrically similar • Similarity parameters (Re, M) are same Aerodynamics 2015 fall - 9 - Introduction to Aerodynamics < 1.8. -
Aerodynamic Force Measurement on an Icing Airfoil
AerE 344: Undergraduate Aerodynamics and Propulsion Laboratory Lab Instructions Lab #13: Aerodynamic Force Measurement on an Icing Airfoil Instructor: Dr. Hui Hu Department of Aerospace Engineering Iowa State University Office: Room 2251, Howe Hall Tel:515-294-0094 Email:[email protected] Lab #13: Aerodynamic Force Measurement on an Icing Airfoil Objective: The objective of this lab is to measure the aerodynamic forces acting on an airfoil in a wind tunnel using a direct force balance. The forces will be measured on an airfoil before and after the accretion of ice to illustrate the effect of icing on the performance of aerodynamic bodies. The experiment components: The experiments will be performed in the ISU-UTAS Icing Research Tunnel, a closed-circuit refrigerated wind tunnel located in the Aerospace Engineering Department of Iowa State University. The tunnel has a test section with a 16 in 16 in cross section and all the walls of the test section optically transparent. The wind tunnel has a contraction section upstream the test section with a spray system that produces water droplets with the 10–100 um mean droplet diameters and water mass concentrations of 0–10 g/m3. The tunnel is refrigerated by a Vilter 340 system capable of achieving operating air temperatures below -20 C. The freestream velocity of the tunnel can set up to ~50 m/s. Figure 1 shows the calibration information for the tunnel, relating the freestream velocity to the motor frequency setting. Figure 1. Wind tunnel airspeed versus motor frequency setting. The airfoil model for the lab is a finite wing NACA 0012 with a chord length of 6 inches and a span of 14.5 inches. -
Active Control of Flow Over an Oscillating NACA 0012 Airfoil
Active Control of Flow over an Oscillating NACA 0012 Airfoil Dissertation Presented in Partial Fulfillment of the Requirements for the Degree Doctor of Philosophy in the Graduate School of The Ohio State University By David Armando Castañeda Vergara, M.S., B.S. Graduate Program in Aeronautical and Astronautical Engineering The Ohio State University 2020 Dissertation Committee: Dr. Mo Samimy, Advisor Dr. Datta Gaitonde Dr. Jim Gregory Dr. Miguel Visbal Dr. Nathan Webb c Copyright by David Armando Castañeda Vergara 2020 Abstract Dynamic stall (DS) is a time-dependent flow separation and stall phenomenon that occurs due to unsteady motion of a lifting surface. When the motion is sufficiently rapid, the flow can remain attached well beyond the static stall angle of attack. The eventual stall and dynamic stall vortex formation, convection, and shedding processes introduce large unsteady aerodynamic loads (lift, drag, and moment) which are undesirable. Dynamic stall occurs in many applications, including rotorcraft, micro aerial vehicles (MAVs), and wind turbines. This phenomenon typically occurs in rotorcraft applications over the rotor at high forward flight speeds or during maneuvers with high load factors. The primary adverse characteristic of dynamic stall is the onset of high torsional and vibrational loads on the rotor due to the associated unsteady aerodynamic forces. Nanosecond Dielectric Barrier Discharge (NS-DBD) actuators are flow control devices which can excite natural instabilities in the flow. These actuators have demonstrated the ability to delay or mitigate dynamic stall. To study the effect of an NS-DBD actuator on DS, a preliminary proof-of-concept experiment was conducted. This experiment examined the control of DS over a NACA 0015 airfoil; however, the setup had significant limitations. -
In-Flight Camber Transformation of Aircraft Wing Based on Variation in Thickness-To-Chord Ratio
International Journal of Industrial Electronics and Electrical Engineering, ISSN: 2347-6982 Volume-2, Issue-7, July-2014 IN-FLIGHT CAMBER TRANSFORMATION OF AIRCRAFT WING BASED ON VARIATION IN THICKNESS-TO-CHORD RATIO 1PRITAM KUMAR PRATIHARI, 2NAVUDAY SHARMA, 3JAYRAJINAMDAR 1,2,3Amity Institute of Space Science and Technology, Amity University, Noida, 201303, Uttar Pradesh, India. E-mail: [email protected], [email protected], [email protected] Abstract- In general, there is wide gap between subsonic flight and supersonic flight, mainly due to aerodynamic limitations. This paper presents a method for development of next generation aircraft wings whose camber can be transformed in-flight for enabling it to fly efficiently in different regimes of Mach number; say, Subsonic to supersonic and vice-versa. To demonstrate the aerodynamic performance of the subsonic airfoil (C141A) and a supersonic airfoil (NACA 65206), the flow is simulated by using the commercial program GAMBIT and FLUENT. An important aspect is the mechanical and hydraulic transformation mechanism working inside the wing which helps it to change the camber according to the requirement. For this mechanism, a Hydraulic expansion-contraction system is used which works on the principle of Master and Slave Piston-cylinder arrangement. This concept can be practically very important to increase aerodynamic efficiency of the aircraft in different regions of subsonic, transonic and supersonic flights. Moreover, this idea of real-time camber transformation can make an aircraft multipurpose in accordance to Mach regime of flight. I. INTRODUCTION transform the camber of the subsonic airfoil to supersonic airfoil and vice versa in flight by variation Much research on the development of aircraft wings of thickness to chord ratio of the airfoil. -
Supersonic Biplane Design Via Adjoint Method A
SUPERSONIC BIPLANE DESIGN VIA ADJOINT METHOD A DISSERTATION SUBMITTED TO THE DEPARTMENT OF AERONAUTICS AND ASTRONAUTICS AND THE COMMITTEE ON GRADUATE STUDIES OF STANFORD UNIVERSITY IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF DOCTOR OF PHILOSOPHY Rui Hu May 2009 c Copyright by Rui Hu 2009 All Rights Reserved ii I certify that I have read this dissertation and that, in my opinion, it is fully adequate in scope and quality as a dissertation for the degree of Doctor of Philosophy. (Antony Jameson) Principal Adviser I certify that I have read this dissertation and that, in my opinion, it is fully adequate in scope and quality as a dissertation for the degree of Doctor of Philosophy. (Robert W. MacCormack) I certify that I have read this dissertation and that, in my opinion, it is fully adequate in scope and quality as a dissertation for the degree of Doctor of Philosophy. (Gianluca Iaccarino) Approved for the University Committee on Graduate Studies. iii Abstract In developing the next generation supersonic transport airplane, two major challenges must be resolved. The fuel efficiency must be significantly improved, and the sonic boom propagating to the ground must be dramatically reduced. Both of these objec- tives can be achieved by reducing the shockwaves formed in supersonic flight. The Busemann biplane is famous for using favorable shockwave interaction to achieve nearly shock-free supersonic flight at its design Mach number. Its performance at off-design Mach numbers, however, can be very poor. This dissertation studies the performance of supersonic biplane airfoils at design and off-design conditions. -
Supersonic Thin Airfoil Theory Aa200b Lecture 5 January 20, 2005
Supersonic Thin Airfoil Theory AA200b Lecture 5 January 20, 2005 AA200b - Applied Aerodynamics II 1 AA200b - Applied Aerodynamics II Lecture 5 Foundations of Supersonic Thin Airfoil Theory The equations that govern the steady subsonic and supersonic flow of an inviscid fluid around thin bodies can be shown to be very similar in appearance to Laplace’s equation @2Á @2Á + = 0; (1) @x2 @y2 which we have used in previous lectures for incompressible flow. In fact, the equations for a free stream Mach number different from zero are given by @2Á @2Á (1 ¡ M 2 ) + = 0: (2) 1 @x2 @y2 Notice that although Eqns. 1 and 2 look very similar, a fundamental change in behavior occurs as the free stream Mach Number, M1, increases past 2 AA200b - Applied Aerodynamics II Lecture 5 the value of 1: the PDE goes from elliptic to hyperbolic and it takes on a wavelike behavior. Eqn. 2 governs the behavior of a supersonic flow where only small disturbances have been introduced. You should already be familiar with a number of results for two- dimensional, isentropic, supersonic flow such as the Prandtl-Meyer function (weak reversible turning analysis, both positive and negative). These notes expand on that knowledge an apply it to the computation of forces and moments on supersonic airfoils. The main difference when compared to incompressible thin airfoil theory is the appearance of drag (in an inviscid flow). The source of this drag is the radiation to infinity of pressure waves created by disturbances at the surface of the airfoil. We will call this drag wave drag. -
Probabilistic Aerothermal Design of Compressor Airfoils Victor E. Garzon
Probabilistic Aerothermal Design of Compressor Airfoils by Victor E. Garzon B.S., University of Kentucky (1994) M.S., University of Kentucky (1998) Submitted to the Department of Aeronautics and Astronautics in partial fulfillment of the requirements for the degree of Doctor of Philosophy at the MASSACHUSETTS INSTITUTE OF TECHNOLOGY February 2003 c Massachusetts Institute of Technology 2003. All rights reserved. Author Department of Aeronautics and Astronautics 11 December 2002 Certified by David L. Darmofal Associate Professor of Aeronautics and Astronautics Thesis Supervisor Certified by Mark Drela Professor of Aeronautics and Astronautics Certified by Edward M. Greitzer H. N. Professor of Aeronautics and Astronautics Certified by Ian A. Waitz Professor of Aeronautics and Astronautics Accepted by Edward M. Greitzer H. N. Slater Professor of Aeronautics and Astronautics Chair, Committee on Graduate Students Probabilistic Aerothermal Design of Compressor Airfoils by Victor E. Garzon Submitted to the Department of Aeronautics and Astronautics on 11 December 2002, in partial fulfillment of the requirements for the degree of Doctor of Philosophy Abstract Despite the generally accepted notion that geometric variability is undesirable in turbo- machinery airfoils, little is known in detail about its impact on aerothermal compressor performance. In this work, statistical and probabilistic techniques were used to assess the impact of geometric and operating condition uncertainty on axial compressor performance. High-fidelity models of geometric variability were constructed from surface measurements of existing hardware using principal component analysis (PCA). A quasi-two-dimensional cascade analysis code, at the core of a parallel probabilistic analysis framework, was used to assess the impact of uncertainty on aerodynamic performance of compressor rotor airfoils. -
List of Symbols
List of Symbols a atmosphere speed of sound a exponent in approximate thrust formula ac aerodynamic center a acceleration vector a0 airfoil angle of attack for zero lift A aspect ratio A system matrix A aerodynamic force vector b span b exponent in approximate SFC formula c chord cd airfoil drag coefficient cl airfoil lift coefficient clα airfoil lift curve slope cmac airfoil pitching moment about the aerodynamic center cr root chord ct tip chord c¯ mean aerodynamic chord C specfic fuel consumption Cc corrected specfic fuel consumption CD drag coefficient CDf friction drag coefficient CDi induced drag coefficient CDw wave drag coefficient CD0 zero-lift drag coefficient Cf skin friction coefficient CF compressibility factor CL lift coefficient CLα lift curve slope CLmax maximum lift coefficient Cmac pitching moment about the aerodynamic center CT nondimensional thrust T Cm nondimensional thrust moment CW nondimensional weight d diameter det determinant D drag e Oswald’s efficiency factor E origin of ground axes system E aerodynamic efficiency or lift to drag ratio EO position vector f flap f factor f equivalent parasite area F distance factor FS stick force F force vector F F form factor g acceleration of gravity g acceleration of gravity vector gs acceleration of gravity at sea level g1 function in Mach number for drag divergence g2 function in Mach number for drag divergence H elevator hinge moment G time factor G elevator gearing h altitude above sea level ht altitude of the tropopause hH height of HT ac above wingc ¯ h˙ rate of climb 2 i unit vector iH horizontal -
Aerodynamic Influence Coefficient Comptuations Using Euler Navier
AIAA JOURNAL Vol. 37, No. 11, November 1999 Aerodynamic Influence Coefficient Computations Using Euler/Navier-Stokes Equations on Parallel Computers Chansup Byun,* Mehrdad Farhangnia, Gurd an Guruswamy. fu P * NASA Ames Research Center, Moffett Field, California 94035-1000 efficienn A t procedur computo et e aerodynamic influence coefficients (AICs), using high-fidelity flow equations suc Euler/Navier-Stokes ha s equations presenteds ,i AICe computee .Th sar perturbiny db g structures using mode shapes. The procedure is developed on a multiple-instruction, multiple-data parallel computer. In addition to discipline parallelizatio coarse-graid nan n parallelizatio floe th w f ndomaino , embarrassingly parallel implemen- tation of ENSAERO code demonstrates linear speedup for a large number of processors. Demonstration of the AIC computatio statir nfo c aeroelasticity analysi arromads sn i a n weo wing-body configuration. Validatioe th f no current procedure is made on a straight wing with arc-airfoil at a subsonic region. The present flutter speed and frequency of the wing show excellent agreement with those results obtained by experiment and NASTRAN. The demonstrated linear scalability for multiple concurrent analyses shows that the three-level parallelism in the code is well suited for the computation of the AICs. Introduction transonic flows over fighter wings undergoing unsteady motions at ODERN design requirement aircrafr sfo t push current tech- small to moderately large angles of attack.3'5 The code has been M nologie sdesige useth n di n proces theio t s r limit somer so - extended to simulate unsteady flows over rigid wing and wing-body times require more advanced technologie meeo st requirementse tth . -
Supersonic Channel Airfoils for Reduced Drag
AIAA JOURNAL Vol. 38, No. 3, March 2000 Supersonic Channel Airfoils for Reduced Drag † ‡ Stephen M. Ruf� n,¤ Anurag Gupta, and David Marshall Georgia Institute of Technology, Atlanta, Georgia 30332-0150 A proof-of-concept study is performed for a supersonic channel-airfoil concept, which can be applied to the leading edges of wings, tails, � ns, struts, and other appendages of aircraft, atmospheric entry vehicles, and missiles in supersonic � ight. It is designed to be bene� cial at conditions in which the leading edge is signi� cantly blunted and the Mach number normal to the leading edge is supersonic. The supersonic channel-airfoil concept is found to result in signi� cantly reduced wave drag and total drag (including skin-friction drag) and signi� cantly increased lift/drag although maximum heat-transfer rate was increased for the geometries tested. Introduction In the present paper a preliminary investigation of a drag-reduc VARIETY of supersonicand hypersonicvehiclesare currently tion concept for supersonicairfoils and wings is conducted.A wide A being studied for commercial and military applications. The range of geometric parameters and supersonic � ight conditions are high-speed civil transport (HSCT) aircraft is designed to cruise at considered so that the aerothermodynamic performance of airfoils approximately M = 2.4 and seeks to overcome economic and en employing the present concept are characterized. vironmental barriers1 that have limited the success of previous su personic commercial concepts. Other supersonic � ight vehicles of Overview of Concept signi� cant interest include single-stage-to-orbit (SSTO) and multi Linearizedsupersonictheoryindicatesthatfor an airfoilof a given stage launch vehicles, tactical and strategic hypersonic and super thickness the shape that gives minimum zero-lift bluntness drag is sonic missiles, hypersonic cruise aircraft, and planetary entry ve the sharp diamond airfoil. -
Aerodynamic Forces on a Stationary and Oscillating Circular Cylinder at High Reynolds Numbers
N ASA TECHNICA L REPORT 0 0 M I w w c 4 m 4 z AERODYNAMIC FORCES ON A STATIONARY AND OSCILLATING CIRCULAR CYLINDER AT HIGH REYNOLDS NUMBERS bY George W.Jones, Jr. Lungley Reseurch Center Joseph J. Cincotta The Martin Company and Robert W. WuZker George C. Mdrslbu ZZ Space FZight Center NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. FEBRUARY 1969 TECH LIBRARY KAFB, NW I llllll11111 lllll I11111I llll11111lll II Ill 0068432 AERODYNAMIC FORCES ON A STATIONARY AND OSCILLATING CIRCULAR CYLINDER AT HIGH REYNOLDS NUMBERS By George W. Jones, Jr. Langley Research Center Langley Station, Hampton, Va. Joseph J. Cincotta The Martin Company Baltimore, Md. and Robert W. Walker George C. Marshall Space Flight Center Huntsville, Ala. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - CFSTI price $3.00 CONTENTS Page SUMMARY ....................................... 1 INTRODUCTION .................................... 2 SYMBOLS ....................................... 3 APPARATUS AND TESTS ............................... 7 Test Facility ..................................... 7 Model ........................................ 7 Instrumentation and Dah-Reduction Procedures .................. 10 Tests ......................................... 11 RESULTS AND DISCUSSION .............................. 12 Static Measurements ................................ 12 Static pressures .................................. 12 Dragdata ....................................