DOCSLIB.ORG

And High Power, the Dynamic Pressure in the Shaded Area Can Be Much

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
And High Power, the Dynamic Pressure in the Shaded Area Can Be Much NAVWEPS 00-801-80 BASIC AERODYNAMICS and high power, the dynamic pressure in the net lift produced by the airfoil is difference shaded area can be much greater than the free between the lifts on the upper and lower sur- stream and this causes considerably greater faces. The point along the chord where the lift than at zero thrust. At high power con- distributed lift is effectively concentrated is ditions the induced flow also causes an effect termed the "center of pressure, c.p. - The similar to boundary layer control and increases center of pressure is essentially the ''center of the maximum lift angle of attack. The typical gravity" of the distributed lift pressure and four-engine propeller driven airplane may have the location of the c.p. is a function of camber 60 to 80 percent of the wing area affected by and section lift coefficient. the induced flow and power effects on stall Another aerodynamic reference point is the speeds may be considerable. Also, the lift of -aerodynamic center, a.c. - The aerodynamic the airplane at a given angle of attack and air- center is defined as the point along the chord speed will be greatly affected. Suppose the where all changes in lift effectively take place. airplane shown is in the process of landing To visualize the existence of such a point, flare from a power-on approach. If there is notice the change in pressure distribution with a sharp, sudden reduction of power, the air- angle of attack for the symmetrical airfoil plane may drop suddenly because of the reduced of figure 1.21. When at zero lift, the upper lift. and lower surface lifts are equal and located The typical jet aircraft does not experience at the same point. With an increase in angle the induced flow velocities encountered in of attack, the upper surface lift increases while propeller driven airplanes, thus the only sig- the lower surface lift decreases. The change nificant factor is the vertical component of of lift has taken place with no change in the thrust. Since this vertical component con- center of pressure—a characteristic of sym- tributes to supporting the airplane, less aero- metrical airfoils. dynamic lift is required to hold the airplane Next, consider the cambered airfoil of in flight. If the thrust is small and the thrust figure 1.21 at zero lift. To produce zero lift, inclination is slight at maximum lift angle, the upper and lower surface lifts must be equal. only negligible changes in stall speed will re- One difference noted from the symmetrical air- sult. On the other hand, if the thrust is very foil is that the upper and lower surface lifts are great and is given a large inclination at maxi- not opposite one another. While no net lift mum lift angle, the effect on stall speed can exists on the airfoil, the couple produced by be very large. One important relationship the upper and lower surface lifts creates a nose remains since there is very little induced flow down moment. As the angle of attack is in- from the jet, the angle of attack at stall is creased, the upper surface lift increases while essentially the same power-on or power-off. the lower surface lift decreases. While a change in lift has taken place, no change in DEVELOPMENT OF AERODYNAMIC PITCHING MOMENTS moment takes place about the point where the lift change occurs. Since the moment The distribution of pressure over a surface about the aerodynamic center is the product is the source of the aerodynamic moments as of a force (lift at the c.p.) and a lever arm well as the aerodynamic forces. A typical example of this fact is the pressure distribution (distance from c.p. to a.c.), an increase in lift acting on the cambered airfoil of figure 1.21. moves the center of pressure toward the aero- The upper surface has pressures distributed dynamic center. which produce the upper surface lift; the lower It should be noted that the symmetrical air- surface has pressures distributed which pro- foil at zero lift has no pitching moment about duce the lower surface lift. Of course, the the aerodynamic center because the upper and 47 Aircraft Technical Book Company (800) 780-4115 (970)-887-2207 http://www.ACTechBooks.com NAVWEPS 00--80T-80 BASIC AERODYNAMICS CAMBERED AIRFOIL UPPER DEVELOPING POSITIVE SURFACE LIFT LIFT NET LIFT C. P. -LOWER SURFACE LIFT SYMMETRICAL AIRFOIL CAMBERED AIRFOIL AT ZERO LIFT AT ZERO LIFT UPPER SURFACE UPPER SURFACE LIFT LIFT .. 41/111W .40er LOWER SURFACE 41! IOW LIFT LOWER SURFACE LIFT SYMMETRICAL AIRFOIL CAMBERED AIRFOIL AT POSITIVE LIFT AT POSITIVE LIFT I UPPER SURFACE LIFT 1-4-UPPER SURFACE LIFT -way w 111111rmr. LOWER SURFACE LIFT LOWER SURFACE LIFT 4--CHANGE IN LIFT -4-CHANGE IN LIFT f t PITCHING MOMENT a.c. a.c. Figure 1.21. Development of Pitching Moments 48 Aircraft Technical Book Company (800) 780-4115 (970)-887-2207 http://www.ACTechBooks.com NAVWEPS 00-801-80 BASIC AERODYNAMICS lower surface lifts act along the same vertical c,,,,,.e. versus lift coefficient for several repre- line. An increase in lift on the symmetrical sentative sections. The sign convention ap- airfoil produces no change in this situation and plied to moment coefficients is that the nose-up the center of pressure remains fixed at the aero- moment is positive. dynamic center. The NACA 0009 airfoil is a symmetrical sec- The location of the aerodynamic center of an tion of 9 percent maximum thickness. Since airfoil is not affected by camber, thickness, and the mean line of this airfoil has no camber, angle of attack. In fact, two-dimensional in- the coefficient of moment about the aerody- compressible airfoil theory will predict the namic center is zero, i.e., the c.p. is at the a.c. aerodynamic center at the 25 percent chord point The departure from zero occurs only as the for any airfoil regardless of camber, thickness, airfoil approaches maximum lift and the stall and angle of attack. Actual airfoils, which produces a moment change in the negative are subject to real fluid flow, may not have the (nose-down) direction. The NACA 4412 and lift due to angle of attack concentrated at the 63 1 -412 sections have noticeable positive cam- exact 25 percent chord point. However, the ber which cause relatively large moments about actual location of the aerodynamic center for the aerodynamic center. Notice that for each various sections is rarely forward of 23 percent section shown in figure 1.22, the cm..e. is constant or aft of 27 percent chord point. for all lift coefficients less than clmax. The moment about the aerodynamic center The NACA 23012 airfoil is a very efficient has its source in the relative pressure distribu- conventional section which has been used on tion and requires application of the coefficient many airplanes. One of the features of the form of expression for proper evaluation. The section is a relatively high coma, with only a moment about the aerodynamic center is ex- small cia.e. The pitching moment coefficients pressed by the following equation: for this section are shown on figure 1.22 along with the effect of various type flaps added to = qSc the basic section. Large amounts of camber where applied well aft on the chord cause large nega- tive moment coefficients. This fact is illus- Ma.c. = moment about the aerodynamic center, trated by the large negative moment coeffi- a.c., ft.-lbs. cients produced by the 30° deflection of a 25 percent chord flap. Cma.c. = coefficient of moment about the a.c. The cn, a, is a quantity determined by the q= dynamic pressure, psf shape of the mean-camber line. Symmetrical airfoils have zero cma and the c.p. remains at S= wing area, sq ft. the a.c. in unstalled flight. The airfoil with positive camber will have a negative c,„„, c= chord, ft. which means the c.p. is behind the a.c. Since the cma.c. is constant in unstalled flight a certain The moment coefficient used in this equation is relationship between lift coefficient and center the dimensionless ratio of the moment pressure of pressure can be evolved. An example of to dynamic pressure moment and is a function this relationship is shown in figure 1.22 for the NACA 63 1-412 airfoil by a plot of c.p. versus Ma.c. Cm c1 — qScqSc . Note that at low lift coefficients the center of pressure is well aft—even past the trailing of the shape of the airfoil mean camber line. edge—and an increase in c 1 moves the c.p. for- Figure 1.22 shows the moment coefficient, ward toward the a.c. The c.p. approaches the 49 Revised January 1965 Aircraft Technical Book Company (800) 780-4115 (970)-887-2207 http://www.ACTechBooks.com NAVWEPS 00-801-80 BASIC AERODYNAMICS +0.1 SECTION LIFT COEFFICIENT, C1 .2 .4 .6 .8 1.0 1.2 1.4 16 1.8 0 i NACA 23012 NACA 0009 NACA 31-412 1-7z L.LJ NACA 4412 5 —0.1 s7C--) 0 i---z 25% NACA 23012 WITH SPLIT FLAP AT 30° 0 —0.2 2 0 ►s= 25°4 NACA 23012 WITH PLAIN FLAP AT 30° —0.3 2 5°41 1 1 NACA 23012 WITH SLOTTED FLAP AT 30° —0.4 ehq CP POSITION FOR NACA 631-412 z 2.0 SECTION Cmac = —.075 w c.) 1 .8 1.6 'AMAX 0 1.4 Li 1.2 z 1.0 0 .8 (.3 .6 .4 .2 t-- 0 10 20 30 40 50 60 70 80 90 100 110 CP POSITION PERCENT CHORD AFT OF LEADING EDGE Figure 1.22.
Load more

Recommended publications
  • 6. Aerodynamic-Center Considerations of Wings I
    m 6. AERODYNAMIC-CENTER CONSIDERATIONS OF WINGS I AND WING-BODY COMBINATIONS By John E. Lsmar and William J. Afford, Jr. NASA Langley Research Center SUMMARY Aerodynamic-center variations with Mach number are considered for wings of different planform. The normalizing parameter used is the square root of the wing area, which provides a more meaningful basis for comparing the aerodynamic-center shifts than does the mean geometric chord. The theoretical methods used are shown to be adequate for predicting typical aerodynamlc-center shifts, and ways of minimizing the shifts for both fixed and variable-sweep wings are presented. INTRODUCTION In the design of supersonic aircraft, a detailed knowledge of the aerodynamlc-center position is important in order to minimize trim drag, maximize load-factor capability, and provide acceptable handling qualities. One of the principal contributions to the aerodynamic-center movement is the well-known change in load distributionwithMach number in going from subsonic to supersonic speeds. In addition, large aerodynamic-center var- iations are quite often associated with variable-geometry features such as variable wing sweep. The purpose of this paper is to review the choice of normalizing param- eters and the effects of Mach number on the aerodynamic-center movement of rigid wing-body combinations at low lift. For fixed wings the effects of both conventional and composite planforms on the aerodynamic-center shift are presented, and for variable-sweepwings the characteristic movements of aerodynamlc-center position wlth pivot location and with variable-geometry apex are discussed. Since systematic experimental investigations of the effects of planform on the aerodynamic-center movement with Mach number are still limited, the approach followed herein is to establish the validity of the computative processes by illustrative comparison wlth experiment and then to rely on theory to show the systematic variations.
    [Show full text]
  • CHAPTER TWO - Static Aeroelasticity – Unswept Wing Structural Loads and Performance 21 2.1 Background
    Static aeroelasticity – structural loads and performance CHAPTER TWO - Static Aeroelasticity – Unswept wing structural loads and performance 21 2.1 Background ........................................................................................................................... 21 2.1.2 Scope and purpose ....................................................................................................................... 21 2.1.2 The structures enterprise and its relation to aeroelasticity ............................................................ 22 2.1.3 The evolution of aircraft wing structures-form follows function ................................................ 24 2.2 Analytical modeling............................................................................................................... 30 2.2.1 The typical section, the flying door and Rayleigh-Ritz idealizations ................................................ 31 2.2.2 – Functional diagrams and operators – modeling the aeroelastic feedback process ....................... 33 2.3 Matrix structural analysis – stiffness matrices and strain energy .......................................... 34 2.4 An example - Construction of a structural stiffness matrix – the shear center concept ........ 38 2.5 Subsonic aerodynamics - fundamentals ................................................................................ 40 2.5.1 Reference points – the center of pressure..................................................................................... 44 2.5.2 A different
    [Show full text]
  • An Aerodynamic Comparative Analysis of Airfoils for Low-Speed Aircrafts
    Published by : International Journal of Engineering Research & Technology (IJERT) http://www.ijert.org ISSN: 2278-0181 Vol. 5 Issue 11, November-2016 An Aerodynamic Comparative Analysis of Airfoils for Low-Speed Aircrafts Sumit Sharma B.E. (Mechanical) MPCCET, Bhilai, C.G., India M.E. (Thermal) SSCET, SSGI, Bhilai, C.G., India Abstract - This paper presents investigation on airfoil S1223, NACA4412 at specified boundary conditions, A S819, S8037 and S1223 RTL at low speeds to find out the most comparison is done between results obtained from CFD suitable airfoil design to be used in the low speed aircrafts. simulation in ANSYS and results obtained from an The method used for analysis of airfoils is Computational experiment done in the wind tunnel. The comparative Fluid Dynamics (CFD). The numerical simulation of low result shows that, the result from experiment and from speed and high-lift airfoil has been done using ANSYS- FLUENT (version 16.2) to obtain drag coefficient, lift CFD simulation showing close agreement between these coefficient, coefficient of moment and Lift-to-Drag ratio over two methods. So, the CFD simulation using ANSYS- the airfoils for the comparative analysis of airfoils. The S1223 FLUENT can be used as an relative alternative to RTL airfoil has been chosen as the most suitable design for experimental method in determining drag and lift. the specified boundary conditions and the Mach number from 0.10 to 0.30. Jasminder Singh et al [2] (June- 2015), Analysis is done on NACA 4412 and Seling 1223 airfoils using Computational Key words: CFD, Lift, Drag, Pitching, Lift-to-Drag ratio.
    [Show full text]
  • Lifting Surface Lift and Pitch Considerations
    Stability and Control Some Characteristics of Lifting Surfaces, and Pitch-Moments The lifting surfaces of a vehicle generally include the wings, the horizontal and vertical tail, and other surfaces such as canards, winglets, etc. What they all have in common is that there purpose is to create an aerodynamic force and moment. Whether the surface is infinite (2-D) or finite (3-D), we can assign to it a force and moment. Usually we designate a 2-D lifting surface as an airfoil, and a 3-D lifting surface as a wing (or tail, or winglet, of whatever). Here all lifting surfaces will be assigned the name - wing. Generally the information concerning forces and moments on a wing are determined from wind tunnel tests or from computer codes. Unlike a force, a moment must be referred to a specific reference point. Often time the information given is not referenced to the point of your interest. Consequently we would like to be able to convert results referenced to one point to equivalent information with respect to a different reference point. We will later use these same ideas applied to a complete aircraft. Consider two identical wings mounted in a wind tunnel. One wing is supported at point A, and the other at point B. At each of these support points there is an electronic “balance” used to measure the forces and moments acting on the wing at those respective points. For simplicity the wings will be represented as flat plates. Si nce the wings are identical we can expect that the aerodynamic forces will be the same.
    [Show full text]
  • Preliminary Design and CFD Analysis of a Fire Surveillance Unmanned Aerial Vehicle Luis E
    TFAWS-08-1034 Preliminary Design and CFD Analysis of a Fire Surveillance Unmanned Aerial Vehicle Luis E. Casas1, Jon M. Hall1, Sean A. Montgomery1, Hiren G. Patel1, Sanjeev S. Samra1, Joe Si Tou1, Omar Quijano2, Nikos J. Mourtos3, Periklis P. Papadopoulos3 Aerospace Engineering San Jose State University One Washington Square San Jose, CA, 95192-0087 The Spartan Phoenix is an Unmanned Aerial Vehicle (UAV) designed for fire surveillance. It was inspired by the summer 2007 wildfires in Greece and California that caused billions of dollars in damage and claimed hundreds of lives. The Spartan Phoenix will map the perimeter of wildfires and provide real time feedback as to fire location and growth pattern. This paper presents the preliminary design of the UAV along with a CFD study of the horizontal and vertical stabilizers of the aircraft. Both surfaces use a NACA 0012 airfoil. First, the flow field over the airfoil was constructed using a CFD-GEOM grid. Next the grid was imported into CFD-FASTRAN and CFD-ACE to simulate a subsonic flow through the grid. The simulation was run using the Spartan Phoenix freestream conditions of 45 m/s airspeed at 00 angle-of-attack, and 20 m/s airspeed at 80 angle-of-attack, with pressure and temperature at 1 atm and 300 K respectively. The generated CFD solutions compared favorably to published data, solutions generated by the Institute of Computational Fluid Dynamics (iCFD), as well as results obtained from airfoil analysis using Sub2D, a potential flow simulation software. Nomenclature AR = aspect ratio = angle of attack CL = lift coefficient CL,min = lift coefficient for minimum drag CD = drag coefficient CDi = induced drag coefficient CD0 = zero lift drag coefficient CG = center of gravity c = chord length e = Oswald efficiency factor L/D = lift-to-drag ratio UAV = Unmanned Aerial Vehicle I.
    [Show full text]
  • Introduction to Aircraft Stability and Control Course Notes for M&AE 5070
    Introduction to Aircraft Stability and Control Course Notes for M&AE 5070 David A. Caughey Sibley School of Mechanical & Aerospace Engineering Cornell University Ithaca, New York 14853-7501 2011 2 Contents 1 Introduction to Flight Dynamics 1 1.1 Introduction....................................... 1 1.2 Nomenclature........................................ 3 1.2.1 Implications of Vehicle Symmetry . 4 1.2.2 AerodynamicControls .............................. 5 1.2.3 Force and Moment Coefficients . 5 1.2.4 Atmospheric Properties . 6 2 Aerodynamic Background 11 2.1 Introduction....................................... 11 2.2 Lifting surface geometry and nomenclature . 12 2.2.1 Geometric properties of trapezoidal wings . 13 2.3 Aerodynamic properties of airfoils . ..... 14 2.4 Aerodynamic properties of finite wings . 17 2.5 Fuselage contribution to pitch stiffness . 19 2.6 Wing-tail interference . 20 2.7 ControlSurfaces ..................................... 20 3 Static Longitudinal Stability and Control 25 3.1 ControlFixedStability.............................. ..... 25 v vi CONTENTS 3.2 Static Longitudinal Control . 28 3.2.1 Longitudinal Maneuvers – the Pull-up . 29 3.3 Control Surface Hinge Moments . 33 3.3.1 Control Surface Hinge Moments . 33 3.3.2 Control free Neutral Point . 35 3.3.3 TrimTabs...................................... 36 3.3.4 ControlForceforTrim. 37 3.3.5 Control-force for Maneuver . 39 3.4 Forward and Aft Limits of C.G. Position . ......... 41 4 Dynamical Equations for Flight Vehicles 45 4.1 BasicEquationsofMotion. ..... 45 4.1.1 ForceEquations .................................. 46 4.1.2 MomentEquations................................. 49 4.2 Linearized Equations of Motion . 50 4.3 Representation of Aerodynamic Forces and Moments . 52 4.3.1 Longitudinal Stability Derivatives . 54 4.3.2 Lateral/Directional Stability Derivatives .
    [Show full text]
  • Longitudinal Static Stability
    Longitudinal Static Stability Some definitions M pitching moment without dimensions C m 1 (so without influence of ρ, V and S) V2 Sc 2 it is a ‘shape’ parameter which varies with the angle of attack. Note the chord c in the denominator because of the unit Nm! L For the wing+aircraft we use the C L 1 surface area of the wing S! VS2 2 L For the tail we use the surface of the tail: SH ! C H LH 1 VS2 2 H Definition of aerodynamic center of a wing: The aerodynamic center (a.c.) is the point around which the moment does not change when the angle of attack changes. We can therefore use Cmac as a constant moment for all angles of attack. The aerodynamic center usually lies around a quarter chord from the leading edge. 1 Criterium for longitudinal static stability (see also Anderson § 7.5): We will look at the consequences of the position of the center of gravity, the wing and the tail for longitudinal static stability. For stability, we need a negative change of the pitching moment if there is a positive change of the angle of attack (and vice versa), so: CC 0 mm 00 0 Graphically this means Cm(α) has to be descending: For small changes we write: dC m 0 d We also write this as: C 0 m 2 When Cm(α) is descending, the Cm0 has to be positive to have a trim point where Cm = 0 and there is an equilibrium: So two conditions for stability: 1) Cm0 > 0; if lift = 0; pitching moment has to be positive (nose up) dC 2) m 0 ( or C 0 ); pitching moment has to become more negative when d m the angle of attack increases Condition 1 is easy to check.
    [Show full text]
  • Introduction to Aircraft Performance and Static Stability
    16.885J/ESD.35J Aircraft Systems Engineering Introduction to Aircraft Performance and Static Stability Prof. Earll Murman September 18, 2003 Today’s Topics • Specific fuel consumption and Breguet range equation • Transonic aerodynamic considerations • Aircraft Performance – Aircraft turning – Energy analysis – Operating envelope – Deep dive of other performance topics for jet transport aircraft in Lectures 6 and 7 • Aircraft longitudinal static stability Thrust Specific Fuel Consumption (TSFC) lb of fuel burned • Definition: TSFC (lb of thrust delivered)(hour) • Measure of jet engine effectiveness at converting fuel to useable thrust • Includes installation effects such as – bleed air for cabin, electric generator, etc.. – Inlet effects can be included (organizational dependent) • Typical numbers are in range of 0.3 to 0.9. Can be up to 1.5 • Terminology varies with time units used, and it is not all consistent. – TSFC uses hours – “c” is often used for TSFC lb of fuel burned – Another term used is c t (lb of thrust delivered)(sec) Breguet Range Equation • Change in aircraft weight = fuel burned dW ctTdt ct TSPC/3600 T thrust • Solve for dt and multiply by Vf to get ds VfdW VfW dW VfL dW ds Vfdt ctT ctT W ctD W • Set L/D, ct, Vf constant and integrate W 3600 L TO R Vf ln TSFC DW empty Insights from Breguet Range Equation W 3600 L TO R Vf ln TSFC DW empty 3600 represents propulsion effects. Lower TSFC is better TSFC V L represents aerodynamic effect. L/D is aerodynamic efficiency fD V La M L . a is constant above 36,000 ft.
    [Show full text]
  • Static Stability and Control
    CHAPTER 2 Static Stability and Control "lsn't it astonishing that all these secrets have been preserved for so many years just so that we could discover them!" Orville Wright, June 7, 1903 2.1 HISTORICAL PERSPECTIVE By the start of the 20th century, the aeronautical community had solved many of the technical problems necessary for achieving powered flight of a heavier-than-air aircraft. One problem still beyond the grasp of these early investigators was a lack of understanding of the relationship between stability and control as well as the influence of the pilot on the pilot-machine system. Most of the ideas regarding stability and control came from experiments with uncontrolled hand-launched gliders. Through such experiments, it was quickly discovered that for a successful flight the glider had to be inherently stable. Earlier aviation pioneers such as Albert Zahm in the United States, Alphonse Penaud in France, and Frederick Lanchester in England contributed to the notion of stability. Zahm, however, was the first to correctly outline the requirements for static stability in a paper he presented in 1893. In his paper, he analyzed the conditions necessary for obtaining a stable equilibrium for an airplane descending at a constant speed. Figure 2.1 shows a sketch of a glider from Zahm's paper. Zahm concluded that the center of gravity had to be in front of the aerodynamic force and the vehicle would require what he referred to as "longitudinal dihedral" to have a stable equilibrium point. In the terminology of today, he showed that, if the center of gravity was ahead of the wing aerodynamic center, then one would need a reflexed airfoil to be stable at a positive angle of attack.
    [Show full text]
  • Design and Computational Fluid Dynamics Analysis of an Idealized Modern Wingsuit
    Washington University in St. Louis Washington University Open Scholarship Engineering and Applied Science Theses & Dissertations McKelvey School of Engineering Spring 5-19-2017 Design and Computational Fluid Dynamics Analysis of an Idealized Modern Wingsuit Maria E. Ferguson Washington University in St Louis Follow this and additional works at: https://openscholarship.wustl.edu/eng_etds Part of the Aerodynamics and Fluid Mechanics Commons Recommended Citation Ferguson, Maria E., "Design and Computational Fluid Dynamics Analysis of an Idealized Modern Wingsuit" (2017). Engineering and Applied Science Theses & Dissertations. 227. https://openscholarship.wustl.edu/eng_etds/227 This Thesis is brought to you for free and open access by the McKelvey School of Engineering at Washington University Open Scholarship. It has been accepted for inclusion in Engineering and Applied Science Theses & Dissertations by an authorized administrator of Washington University Open Scholarship. For more information, please contact [email protected]. WASHINGTON UNIVERSITY IN ST. LOUIS School of Engineering and Applied Science Department of Mechanical Engineering and Materials Science Thesis Examination Committee: Dr. Ramesh Agarwal, Chair Dr. Qiulin Qu Dr. David Peters Design and Computational Fluid Dynamics Analysis of an Idealized Modern Wingsuit by Maria E. Ferguson A thesis presented to the School of Engineering and Applied Science of Washington University in St. Louis in partial fulfillment of the requirements for the degree of Master of Science May
    [Show full text]
  • Full-Scale Ornithopter
    The Flight Dynamics of a Full-Scale Ornithopter Tahir Rashid A thesis submitted in confomity with the requiiements for the degree of Master of Applied Science Graduate Department of Aerospace Science and Engineering in the University of Toronto O Copyright by Tahir Rashid, 1995 National Library Bibliothèque nationale du Canada Acquisitions and Acquisitions et Bibliographie Services services bibliographiques 395 Wdliion Street 395. nie Wellington OnewaûN K1AW Ottawa ON K1A ON4 Canaada Canada The author has granted a non- L'auteur a accordé une licence non exclusive licence allowing the exclusive permettant a la National Library of Canada to Bibliothèque nationale du Canada de reproduce, loan, distribute or sel1 reproduire, prêter, distribuer ou copies of this thesis in rnicroform, vendre des copies de cette thèse sous paper or electronic formats. la forme de microfiche/fiim, de reproduction sur papier ou sur format électronique. The author retauis ownership of the L'auteur conserve la propriéte du copyright in this thesis. Neither the droit d'auteur qui protège cette thèse. thesis nor substantial exûacts ffom it Ni la thèse ni des extraits substantiels may be printed or otherwise de celle-ci ne doivent être imprimés reproduced without the author's ou autrement reproduits sans son permission. autorisation. The Flight ûynamics of a FulCScale Ornithopter Master of Applied Science, 1995 Tahir Rashid Aerospace Science and Engineering University of Toronto This t hesis investigates the non-lineaif light dynamics of a full-scale flapping-wing aircraft (ornithopter). Using simplifying assumptions, the equations of motion were developed for a 2-wing-panel and 3-wing-panel rnodel.
    [Show full text]
  • Module-3 Lecture-13 Stability and Control
    Module-3 Lecture-13 Stability and Control - Discussion on Center of Pressure, Aerodynamic Center and Trim Center of pressure and Aerodynamic center The resultant aerodynamic force and moment acting on body must have the same effect as the distributed load. The resultant moment will depend on where ever the resultant Figure 1: Resultant aerodynamic force and moment force is placed on the body. For example, let x be the coordinate measured along the chord line of an airfoil, from the leading edge towards the trailing edge. The resultant Figure 2: Various coordinate on chord line moment about some arbitrary point on the chord line a distance x from the leading edge be Mx. Then MLE = Mx − xN x C = C − C mLE mx c N Two particular locations along the chord line are of special interest. ◦ xcp ! Center of pressure: The point about which the resultant moment is zero. ◦ xac ! Aerodynamic center: The point about which the change in the resultant moment with respect to the angle of attack is zero. 1 Center of Pressure (xcp) • By definition, Cmcp = 0 • For x = xcp, this gives, x x C = C − C = − xp C mLE mx c N c N x x C cp = − mx c c CN • Hence, the location of xcp at any given angle of attack (α) can be determined from the normal force coefficient and moment coefficient about any point on the airfoil chord line. • In general, xcp may vary significantly with α. Aerodynamic Center (xac) • For x = xac, we have, x x C = C − C = C − ac C mLE mx c N mac c N x x C = C + ac − C mac mx c c N • From definition of aerodynamic center, @C mac = 0 @α @C @C x x @C mac = mx + ac − N = 0 @α @α c c @α • Thus, x x @Cmx ac = − @α c c @CN @α 2 • The location of the aerodynamic centre can be determined from the knowledge of how the normal force coefficient and moment coefficient about any point on the chord line vary with angle of attack.
    [Show full text]