DOCSLIB.ORG

Wind-Tunnel Investigation of the Flight Chasacteristics of a Canard

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Wind-Tunnel Investigation of the Flight Chasacteristics of a Canard ., , -. / ~ ,a ~ i t 1 1 / I -1 I' I , ,Wind-TunnelInvestigation of the Flight Chasacteristics of a Canard Genera1:Aviation Airplane Coniiguration - I Dale R. Satran j . iNA3A-22-2623 ) UIND- TUNKEL INVESTIGkTION OE 138 7 - 1 d 0 3 j IHZ ELIGHT Cii Ait ACiEiiI STILS C F A CANAtiD Fig32 AVI - G AL- A T iUN AIRP LANE CC NEPGUEATION I 60 [NASA) p csci 01A Ullcla s I H1/02 44247 , NASA NASA Technical Paper 2623 1986 Wind-Tunnel Investigation of the Flight Characteristics of a Canard General-Aviation Airplane Configuration Dale R. Satran Langley Research Center Hampton, Virginia National Aeronautics and Space Administration Scientific and Technical Information Branch Summary figuration. A 0.36-scale model was used in the free- flight investigation and was also used to obtain static A 0.36-scale model of a canard general-aviation and dynamic force data to aid in the interpretation airplane with a single pusher propeller and winglets of the free-flight test results. The free-flight tests was tested in the Langley 30- by 60-Foot Wind Tun- were conducted for angles of attack ranging from 7O nel to determine the static and dynamic stability and to 14O. The investigation included tests of the model control and free-flight behavior of the configuration. with high and low canard positions, three center-of- Model variables made testing of the model possible gravity locations, outboard wing leading-edge droop, with the canard in high and low positions, with in- winglets and a center vertical tail, and several roll- creased winglet area, with outboard wing leading- and yaw-control systems. Dynamic force tests were edge droop, with fuselage-mounted vertical fin and also conducted on the 0.36-scale model using the rudder, with enlarged rudders, with dual deflecting forced-oscillation test technique to study the effects rudders, and with ailerons mounted closer to the of two canard vertical positions and using the instal- wing tips. lation of winglets and outboard leading-edge droop The basic model exhibited generally good longitu- on the roll damping of the model. Wool tufts were dinal and lateral stability and control characteristics. installed to aid in flow visualization of the stall pat- The removal of an outboard leading-edge droop de- tern of the wing and canard during the static force graded roll damping and produced lightly damped tests and free-flight tests. roll (wing rock) oscillations. In general, the model exhibited very stable dihedral effect but weak direc- Symbols tional stability. Rudder and aileron control power were sufficiently adequate for control of most flight All longitudinal forces and moments are refer- conditions, but appeared to be relatively weak for enced to the stability-axis system, and all lateral- maneuvering compared with those of more conven- directional forces and moments are referenced to the tionally configured models. body-axis system. The midpoint of the center-of- gravity range is 0.71C ahead of the leading edge of Introduction the wing mean aerodynamic chord. (See fig. 1.) The wing reference area corresponds to that area obtained As part of the NASA general-aviation stall/spin by extending the outboard leading and trailing edges program, advanced aircraft configurations are be- of the wing without leading-edge droop to the fuse- ing investigated that offer unique safety benefits. lage centerline. All dimensional quantities are ex- One such configuration, the R.utan VariEze, uti- pressed in both the International System of Units lizes a high-aspect-ratio canard, a swept-back wing, (SI) and U.S. Customary Units. Measurements were winglets, and a pusher propeller. Full-scale flight made in U.S. Customary Units, and conversion fac- tests of the homebuilt canard aircraft have demon- tors from reference 5 were used to obtain equivalent strated advantages for such a design from the stand- SI dimensions. point of increased stall departure and spin resistance. (See ref. 1.) Several models of this configuration were wing span, cm (in.) tested in different facilities to document the flight drag coefficient, characteristics of the VariEze. Reference 2 contains 9 static wind-tunnel data for a full-scale mode! of the iift coefficient, configuration tested in the Langley 30- by 60-Foot 3 Wind Tunnel. Rotary-balance tests were conducted rolling-moment coefficient, on a 0.22-scale model in the Langley Spin Tunnel, and the results showed the configuration to have in- pitching-moment coefficient, herently good stall departure and spin resistance. Pitching moment qsc (See ref. 3.) In addition to improved safety fea- tures, the configuration has all-composite construc- yawing-moment coefficient, tion, which makes possible a smooth surface finish and, for this particular configuration, the realization effective thrust coefficient at zero angle of attack, Drag (power off)-Drag (power on) of performance gains through large improvements in 9s natural laminar flow. (See ref. 4.) The purpose of this investigation was to use side-force coefficient, Sid:p the free-flight test technique in the Langley 30- by mean aerodynamic chord, cm (in.) 60-Foot Wind Tunnel to study the dynamic stability and control and general flight behavior of the con- frequency of oscillation, Hz moment of inertia about X axis, kg-m2 Abbreviations: (slug-ft2) BL butt line moment of inertia about Y axis, kg-m2 c.g. center of gravity (slug-ft2) FS fuselage station moment of inertia about 2 axis, kg-m2 L.E. leading edge (slug- ft2 ) max maximum incidence angle of canard, positive WL waterline trailing edge down, deg reduced frequency parameter, wb/2V Model and Apparatus roll rate, rad/sec The basic configuration is depicted in a three-view diagram in figure 1, and photographs of the model free-stream dynamic pressure, Pa (psf) are shown in figure 2. The mass and dimensional characteristics are included in table I. The 0.36-scale wing reference area, m2 (ft2) model is representative of the Rutan VariEze, a two- free-stream velocity, m/sec (ft/sec) place, advanced general-aviation airplane. For all tests, the nose gear was retracted. The mass and in- spanwise coordinate, m (ft) ertial characteristics were scaled to correspond to op- angle of attack, deg eration at 1524 m (5000 ft) altitude (standard atmo- sphere). The wing, winglet, and canard of the model angle of sideslip, deg were constructed of balsa wood and fiberglass. The rate of change of sideslip, rad/sec fuselage was made of fiberglass and foam sandwich construction with an internal aluminum structure. increniental rolling-moment coefficient The control surfaces were actuated for free-flight (control deflected - control neutral) tests by electroprieurrlatic servos. The controls con- iricrernental yawing-moment coefficient sisted of a slotted canard flap used as an elevator, ailerons located inboard on the main wing, and rud- (control deflected - control ricutral) ders mounted on the winglets. The basic rudders increinrnt a1 side-force coefficient deflected independently and outward only; that is, (control deflected - control neutral) for a left turn, the trailing edge of the left rudder only would move to the left. The control deflections aileron deflection, positive for left roll, were limited during free-flight tests to f20" for the deg ailerons, f30" for the rudders, and f5" for the el- elevator deflection, positive for trailing evator. Thrust to fly the model was supplied by a edge dow11, deg propeller driven by a turbine-air motor using com- pressed air. flap deflection, positive for trailing edge Static force tests were made with several rud- down, deg der modifications using sheet-metal tabs to simulate dual, split, and enlarged winglet rudders. These rud- rudder deflection, positive for left rudder der modifications are shown in figure 3. The dual and trailing edge left, deg split rudders had an area equivalent to that of the angular frequency, 27r f , rad/sec basic rudders. For dual-rudder control, both rudders deflected simultaneously and in the same direction; Stability derivatives: that is, for a left turn, the trailing edges of both rud- ders moved to the left. For split-rudder control, the inboard arid outboard surfaces of one rudder split and deflected outward from the neutral position, and the rudder oti the opposite winglet remained undeflected. The enlarged rudders operated in the same manner as the basic rudders, but had twice the chord, and extended in height to the tip of the winglets. The hinge line of the rudder was unchanged for all rudder modifications. A center vertical fin and conventional 2 rudder mounted on the fuselage directly ahead of the The static force data were measured at a nominal propeller were also tested. dynamic pressure of 464 Pa (9.7 psf), corresponding Force tests were also conducted with outboard- to a Reynolds number of 0.535 x lo6 based on wing mounted ailerons and with differential elevator de- mean aerodynamic chord or 0.218 x lo6 based on ca- flection for roll control. The outboard ailerons were nard mean aerodynamic chord. The static force tests simulated using sheet-metal tabs mounted on the were conducted over an angle-of-attack range from trailing edge of the wing on the outer 25 percent of -loo to 90' and an angle-of-sideslip range of f15', the span. (See fig. 4.) Landing flaps were simulated although some of the tests were made over reduced in exploratory force tests by deflecting the ailerons angle-of-attack ranges. Static sideslip derivatives symmetrically. were determined from 3x5' sideslip angles. Wind- Tests were conducted with the canard mounted on tunnel flow-angularity corrections were applied to all the top of the fuselage, as the basic location, and with data based on wind-tunnel flow surveys. Because the the canard mounted on the bottom of the fuselage, size of the model was small relative to that of the test as an alternative location.
Load more

Recommended publications
  • A Numerical Approach for Implementing Air Intakes in a Canard Type Aircraft for Engine Cooling Purposes
    https://doi.org/10.1590/jatm.v13.1192 ORIGINAL PAPER A Numerical Approach for Implementing Air Intakes in a Canard Type Aircraft for Engine Cooling Purposes Odenir de Almeida1,* , Pedro Correa Souza1 , Erick Cunha2 1.Universidade Federal de Uberlândia – Faculdade de Engenharia Mecânica – Centro de Pesquisa em Aerodinâmica Experimental – Uberlândia/MG – Brazil. 2.Fábrica Brasileira de Aeronaves – Uberlândia/MG – Brazil *Corresponding author: [email protected] ABSTRACT This work presents selected results of an unconventional aircraft development campaign. Engine installation at the rear part of the fuselage imposed design constraints for air intakes that should be used for cooling purposes. Trial and error flight tests increased the development cost and time which required a more sophisticated analysis through computational fluid dynamics (CFD) techniques and robust semiempirical approach. The carried-out investigation of the air intakes started with an empirical approach from guidelines for designing NACA and scoops. Numerical studies via computational fluid dynamics were performed with the air intakes installed in the aircraft fuselage. An analysis based on the air intake efficiency, drag and the effect of angle of attack are detailed in this work. Different air intakes designs, such as scoops of different shapes, were evaluated seeking for improved air intake efficiency and low drag while providing enough air for cooling the engine compartment. The results showed that the numerical approach used herein decreased the development cost and time of the aircraft, providing a reasonable low-cost approach and leading to a design selection more easily. Based on the current approach the canard airplane geometry was changed to account for the new selected air intake for engine cooling purposes.
    [Show full text]
  • CANARD.WING LIFT INTERFERENCE RELATED to MANEUVERING AIRCRAFT at SUBSONIC SPEEDS by Blair B
    https://ntrs.nasa.gov/search.jsp?R=19740003706 2020-03-23T12:22:11+00:00Z NASA TECHNICAL NASA TM X-2897 MEMORANDUM CO CN| I X CANARD.WING LIFT INTERFERENCE RELATED TO MANEUVERING AIRCRAFT AT SUBSONIC SPEEDS by Blair B. Gloss and Linwood W. McKmney Langley Research Center Hampton, Va. 23665 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, D. C. • DECEMBER 1973 1.. Report No. 2. Government Accession No. 3. Recipient's Catalog No. NASA TM X-2897 4. Title and Subtitle 5. Report Date CANARD-WING LIFT INTERFERENCE RELATED TO December 1973 MANEUVERING AIRCRAFT AT SUBSONIC SPEEDS 6. Performing Organization Code 7. Author(s) 8. Performing Organization Report No. L-9096 Blair B. Gloss and Linwood W. McKinney 10. Work Unit No. 9. Performing Organization Name and Address • 760-67-01-01 NASA Langley Research Center 11. Contract or Grant No. Hampton, Va. 23665 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technical Memorandum National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington , D . C . 20546 15. Supplementary Notes 16. Abstract An investigation was conducted at Mach numbers of 0.7 and 0.9 to determine the lift interference effect of canard location on wing planforms typical of maneuvering fighter con- figurations. The canard had an exposed area of 16.0 percent of the wing reference area and was located in the plane of the wing or in a position 18.5 percent of the wing mean geometric chord above the wing plane. In addition, the canard could be located at two longitudinal stations.
    [Show full text]
  • Actuator Saturation Analysis of a Fly-By-Wire Control System for a Delta-Canard Aircraft
    DEGREE PROJECT IN VEHICLE ENGINEERING, SECOND CYCLE, 30 CREDITS STOCKHOLM, SWEDEN 2020 Actuator Saturation Analysis of a Fly-By-Wire Control System for a Delta-Canard Aircraft ERIK LJUDÉN KTH ROYAL INSTITUTE OF TECHNOLOGY SCHOOL OF ENGINEERING SCIENCES Author Erik Ljudén <[email protected]> School of Engineering Sciences KTH Royal Institute of Technology Place Linköping, Sweden Saab Examiner Ulf Ringertz Stockholm KTH Royal Institute of Technology Supervisor Peter Jason Linköping Saab Abstract Actuator saturation is a well studied subject regarding control theory. However, little research exist regarding aircraft behavior during actuator saturation. This paper aims to identify flight mechanical parameters that can be useful when analyzing actuator saturation. The studied aircraft is an unstable delta-canard aircraft. By varying the aircraft’s center-of- gravity and applying a square wave input in pitch, saturated actuators have been found and investigated closer using moment coefficients as well as other flight mechanical parameters. The studied flight mechanical parameters has proven to be highly relevant when analyzing actuator saturation, and a simple connection between saturated actuators and moment coefficients has been found. One can for example look for sudden changes in the moment coefficients during saturated actuators in order to find potentially dangerous flight cases. In addition, the studied parameters can be used for robustness analysis, but needs to be further investigated. Lastly, the studied pitch square wave input shows no risk of aircraft departure with saturated elevons during flight, provided non-saturated canards, and that the free-stream velocity is high enough to be flyable. i Sammanfattning Styrdonsmättning är ett välstuderat ämne inom kontrollteorin.
    [Show full text]
  • The Design and Development of a Human-Powered
    THE DESIGN AND DEVELOPMENT OF A HUMAN-POWERED AIRPLANE A THESIS Presented to the Faculty of the Graduate Division "by James Marion McAvoy^ Jr. In Partial Fulfillment of the Requirements for the Degree Master of Science in Aerospace Engineering Georgia Institute of Technology June _, 1963 A/ :o TEE DESIGN AND DETERMENT OF A HUMAN-POWERED AIRPLANE Approved: Pate Approved "by Chairman: M(Ly Z7. /q£3 In presenting the dissertation as a partial fulfillment of the requirements for an advanced degree from the Georgia Institute of Technology, I agree that the Library of the Institution shall make it available for inspection and circulation in accordance with its regulations governing materials of this type. I agree that permission to copy from, or to publish from, this dissertation may he granted by the professor under whose direction it was written, or, in his absence, by the dean of the Graduate Division when such copying or publication is solely for scholarly purposes and does not involve potential financial gain. It is under­ stood that any copying from, or publication of, this disser­ tation which involves potential financial gain will not be allowed without written permission. i "J-lW* 11 ACKNOWLEDGMENTS The author wishes to express his most sincere appreciation to Pro­ fessor John J, Harper for acting as thesis advisor, and for his ready ad­ vice at all times. Thanks are due also to Doctor Rohin B. Gray and Doctor Thomas W. Jackson for serving on the reading committee and for their help and ad­ vice . Gratitude is also extended to,all those people who aided in the construction of the MPA and to those who provided moral and physical sup­ port for the project.
    [Show full text]
  • Glider Accidents and Prevention
    Glider Accidents – Statistics & Prevention Larry Suter / John Scott Northern California Soaring Association & Air Sailing Gliderport Glider Accident Summary Data from 2008 – 2013 172 accidents reported to NTSB SSF categorized into three Types 70% 60% 50% 40% Fatal 30% Non Fatal 20% 10% 0% Takeoff Free Flight Landing What fact is painfully obvious? Takeoff 70% 60% 50% 40% Fatal 30% Non Fatal 20% 10% 0% Takeoff Free Flight Landing Approximately 20% of all accidents occur during the takeoff phase Video of a canopy coming open on takeoff. According to the SSF opening canopies and deploying spoilers are more likely to cause a takeoff accident than a rope break or any other type PT3 event. (PT3 = Premature Termination of the Tow) Canopy and spoiler accidents are preventable! They occur because the pilot failed to properly complete their Pre-Takeoff checklist. Low altitude emergency training tends to focus on rope breaks. In reality rope breaks are a small fraction of Takeoff accidents. Deployed Spoilers There is no reason that deployed spoilers should cause an accident. Just close them! Rudder Waggle --- A potentially dangerous signal Two Landmark accidents; same scenario (2006 NV, 2011 MD) Pilots reacted too quickly w/o thinking and pulled release. (Panic? Misinterpretation?) Had insufficient altitude to return to airport. (120’, 200’) ASG Tow Pilot Manual discourages this signal below 1000’ AGL; Rudder deflections and AD coupling. Opening Canopy There is no reason that an opening canopy should cause an accident! Effects of an Opening
    [Show full text]
  • 09 Stability and Control
    Aircraft Design Lecture 9: Stability and Control G. Dimitriadis Introduction to Aircraft Design Stability and Control H Aircraft stability deals with the ability to keep an aircraft in the air in the chosen flight attitude. H Aircraft control deals with the ability to change the flight direction and attitude of an aircraft. H Both these issues must be investigated during the preliminary design process. Introduction to Aircraft Design Design criteria? H Stability and control are not design criteria H In other words, civil aircraft are not designed specifically for stability and control H They are designed for performance. H Once a preliminary design that meets the performance criteria is created, then its stability is assessed and its control is designed. Introduction to Aircraft Design Flight Mechanics H Stability and control are collectively referred to as flight mechanics H The study of the mechanics and dynamics of flight is the means by which : – We can design an airplane to accomplish efficiently a specific task – We can make the task of the pilot easier by ensuring good handling qualities – We can avoid unwanted or unexpected phenomena that can be encountered in flight Introduction to Aircraft Design Aircraft description Flight Control Pilot System Airplane Response Task The pilot has direct control only of the Flight Control System. However, he can tailor his inputs to the FCS by observing the airplane’s response while always keeping an eye on the task at hand. Introduction to Aircraft Design Control Surfaces H Aircraft control
    [Show full text]
  • Design Study of Technology Requirements for High Performance Single-Propeller- Driven Business Airplanes
    3 1176 01346 2362 ! NASA Contractor Report 3863 NASA-CR-386319850007383 Design Study of Technology Requirements for High Performance Single-Propeller- Driven Business Airplanes David L. Kohlman and James Hammer CONTRACT NAS1-16363 JANUARY 1985 N/ A NASA Contractor Report 3863 Design Study of Technology Requirements for High Performance Single-Propeller- Driven Business Airplanes David L. Kohlman and James Hammer Flight Research Laboratory University of Kansas Center for Research, Inc. Lawrence, Kansas Prepared for Langley Research Center under Contract NAS1-16363 N/ A National Aeronautics and Space Administration Scientific and Technical InformationBranch 1985 TABLE OF CONTENTS Page i. INTRODUCTION .............................................. 1 2. NOMENCLATURE .............................................. 5 3. BASELINE CONFIGURATION ANALYSIS ........................... 7 3.1 Description of Baseline Configuration ................ 7 3.2 Effect of Aspect Ratio ............................... ii 3.3 Effect of Wing Loading ............................... 12 3.4 Effect of Wing Natural Laminar Flow .................. 15 3.5 Effect of Fuselage Drag .............................. 19 4. PROPULSION SYSTEM ANALYSIS ................................. 31 4.1 GATE Engine ........................................... 32 4.1.1 Description of Engine .......................... 32 4.1.2 GASP Engine Routine ............................ 33 4.2 Very Advanced Reciprocating Engine (Spark Ignited), (SIR) ............................... 36 4.3 Very Advanced Diesel
    [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]
  • SM11188 Standard Modification Page : 1 of 6 Issue 1 Compiled : S Brown Approved : F Donaldson
    Mod No. SM11188 Standard Modification Page : 1 of 6 Issue 1 Compiled : S Brown Approved : F Donaldson TITLE : Main Wing Trailing Edge Flow Straighteners AIRCRAFT TYPE : Vari-Eze Mod Type: Retro-fit 1. Introduction The Rutan Aircraft Factory approved the (later to be mandatory) installation of leading edge vortilons on all Vari-Ezes in the October 1984 edition of the Canard Pusher ( CP 42 pages 5/6) to improve the swept main wing stall margins, increase visibility over the nose on final approach and enable reduced approach speeds without the risk of low speed wing rock or deep stall. The vortilons generate vortices over the main wing at high angles of attack that reduce lift-losing span-wise flow. There is negligible speed or drag penalty Flow straighteners positioned on the trailing edge of the swept main wing further reduce span wise flow at high angles of attack. It has been demonstrated that the addition of trailing edge flow straighteners enable the wing to produce more lift at a given angle of attack enabling lower approach speeds with no loss of view over the nose. The result is a reduced sink rate for a given airspeed and an increased margin between main wing and canard stall speeds. 2. Parts List Qty Part No. Description Source 2 SM11188-1 Outboard Fence Manufacture 2 SM11188-2 Middle Fence Manufacture 2 SM11188-3 Inboard Fence Manufacture White RTV Silicone Hardware store Flocked Cotton / BID / Aircraft Spruce Epoxy Resin List of related drawings / photos Drawing No. Title / Description Issue SM11188-D1 Drawing of Flow Straighteners dimensions SM11188-D2 Drawing of installation on main wing and fitting detail SM11188-D3 Photograph of Gary Hertzler’s fences installation on N99VE 3.
    [Show full text]
  • Design and Development of a Multi- Mission UAS Through Modular Component Integration and Additive Manufacturing
    Design and Development of a Multi- Mission UAS through Modular Component Integration and Additive Manufacturing A project present to The Faculty of the Department of Aerospace Engineering San Jose State University in partial fulfillment of the requirements for the degree Master of Science in Aerospace Engineering By Kim Lau August, 2018 approved by Dr. Nikos Mourtos Faculty Advisor 2 Table of Contents 1. Introduction ........................................................................................................................5 1.1. Motivation .................................................................................................................5 1.2. Literature Review .....................................................................................................5 1.2.1. UAV Design and Development Trends ...........................................................5 1.2.2. Push for Modularity .........................................................................................7 1.2.3. Additive Manufacturing ...................................................................................8 1.3. Project Proposal ........................................................................................................11 1.4. Methodology .............................................................................................................11 2. Design Process ...................................................................................................................12 2.1. Mission Specification and
    [Show full text]
  • Effects of Controls (Secondary)
    Student Study Guide A Certificate EFFECTS OF CONTROLS (SECONDARY) Aim: To learn about any secondary effects of flight control inputs. We have seen the primary effect of controls… now lets explore any secondary effects… that is… does anything else happen when we pitch, roll or yaw… does one movement have more than the single effect we have seen so far? Secondary Effects: A secondary effect of control can be defined as any effect about one of the 3 main axis of motion which occurs as a result of an initial control input. If we pitch the glider, will it roll or yaw? No… so there is no secondary effect of pitch. How about if we apply aileron and roll the glider… it banks, slips towards the lower wing then yaws towards that lower wing… so yes, there is a secondary effect of roll and it is yaw. And yawing…we apply some rudder…the glider yaws, skids sideways and rolls towards the direction of the yaw. Your instructor will set the glider up in a normal gliding attitude at 45kts and demonstrate these secondary effects so you can see them. To summarise then: Control Axis Primary Effect Secondary Effect Elevato Lateral Glider pitches nose up there is no roll or yaw r and down Aileron Longitudina glider rolls left the glider slips towards the lower wing, then yaws l / right towards the lower wing Rudder Vertical glider yaws left The glider skids then rolls towards the inner wing glider yaws right So What?! Applying elevator to pitch the nose up or down has no secondary effect to consider.
    [Show full text]
  • Aero Dynamic Analysis of Multi Winglets in Light Weight Aircraft
    SSRG International Journal of Mechanical Engineering (SSRG-IJME) – Special Issue ICRTETM March 2019 Aero Dynamic Analysis of Multi Winglets in Light Weight Aircraft J. Mathan#1,L.Ashwin#2, P.Bharath#3,P.Dharani Shankar#4,P.V.Jackson#5 #1Assistant Professor & Mechanical Engineering & KSRIET #2,3,4,5Final Year Student & Mechanical Engineering & KSRIET Tiruchengode,Namakkal(DT),Tamilnadu Abstract An analysis of multi-winglets as a device for of these devices such as winglets [2], tip-sails [3, 4, 5] reducing induced drag in low speed aircraft is and multi-winglets [6] take energy from the spiraling carried out, based on experimental investigations of a air flow in this region to create additional traction. wing-body half model at Re = 4•105. Winglet is a lift This makes possible to achieve expressive gains on augmenting device which is attached at the wing tip efficiency. Whitcomb [2], for example, shows that of an aircraft. A Winglets are used to improve the winglets could increase wing efficiency in 9% and aerodynamic efficiency of an aircraft by lowering the decrease the induced dragin20%. Some devices also formation of an Induced Drag which is caused by the break up the vortices into several parts, each one with wingtip vortices. Numerical studies have been carried less intensity. This facilitates their dispersion, an out to investigate the best aerodynamic performance important factor to decrease the time interval between of a subsonic aircraft wing at various cant angles of takeoff and landings at large airports [7]. A winglets. A baseline and six other different multi comparison of the wingtip devices [1] shows that winglets configurations were tested.
    [Show full text]