Flight Mechanics

Tests in the Ames 40-by 80-foot wind tunnel of the static longitudinal characteristics of the Republic RF -84F were made to determine both the origin and a suitable remedy for a pitch up tendency of the airplane encountered at moderate lift coefficients. The results indicated that the pitch-up at moderate lift coefficients was caused by an abrupt change in downwash at the tail which in turn was traceable presumably to flow conditions associated with the inlet-to-wing leading-edge discontinuity.. Attempts to eliminate this pitch-up characteristic with various fairings and stall-control devices. were not wholly successful. The investigation revealed, however, that significant gains in the performance of the airplane could be achieved in the upper lift range.. Three different configurations consisting of a partial-span modified leading edge combined with one or with two-fenees or a leading-edge extension each delayed the onset of separation to higher lift coefficients and provided large improvements in the stability of the airplane in the upper lift range.

A method is presented whereby incremental pitching moments can be estimated for swept and triangular wings having arbitrary types of trailing-edge high-lift flaps. In the method use is made of span-loading theory together with two-dimensional airfoil data adjusted for the effects of sweep. The method as presented was limited to low speeds and small angles of attack. Application of the method is demonstrated for some 58 cases covering various types of flaps on wings having a wide range of sweep, aspect ratio, and taper ratio. For all wings, swept as well as triangular, a mean deviation from experiment of about 0.02 in incremental pitchingmoment coefficient was found. Two-dimensional-flap data pertinent to the general application of the method are summarized in graphical form.

A method is presented whereby incremental pitching moments canbe estimated forsweptandtriangular wingshaving arbitrary typesof trailing-edge high-lift flaps.In themethod useismadeof span-loading q theory together withtwo-dimensional airfoil dataadjusted fortheeffects of sweep.Themethod as presented waslimited to lowspeeds andsmall q angles of attack. Application of themethcxl is demonstrated forsane58 casescovering various t~es of flapson wingshating a tiderangeof sweep, aspect ratio, andtaperratio.Forallwings, sweptas wellas triangular, a meandeviation frcmexperiment of about0.02in incremental pitchingmomentcoefficient wasfound. Two-dimensional-flap datapertinent to thegeneral application of themethod aresummarized in graphical form.

A study has been made of the application of two-dimensional data and span-loading theory for estimating the local loading characteristics on a swept wing with flaps. Estimated results, including local pressure distributions, span loadings, and the nonlinear local lift characteristics, are compared with similar results measured at large scale on a 45 0 sweptback wing of aspect ratio 6 having a O.4-epan double-elotted flap both with and without a full-span slat. Two-dimensional pressure distributions when corrected for sweep were found to agree closely with the wing pressures for most local sections either on or off the flap. This agreement continued to the higher lift coefficients and even improved near maximum lift where the flap-induced effects became minimized. The Weissinger 7Xl method was found to provide reasonably accurate span lciadings for this awept-wing configuration which had a relatively highly loaded type of flap. Twodimensional lift data, together with span-loading theory, afforded quite accurate estimates of the local nonlinear lift characteristics, includiLg maximum lift of sections outboard of the flap but were inadequate for inboard sections of the wing where the three-dimensional boundary-layer control exercises a dominant effect.

Control allocation compute the commands that are applied to the actuators so that a certain set of forces or moments are generated by the control effectors. In this communication we consider the problem of the design of an airborne system whose function is to solve the problem of allocation of transportation aircraft actuators during manoeuvres. The case of a roll manoeuvre, which may be quite demanding for the wings structure, is more particularly considered. The aim of this supervision function, similar to the load alleviation function, is to control the bending and flexion moments of the wings while performing a manoeuvre which can turn to be extreme. The wings actuators considered are the ailerons, the spoilers and the flaps, but the other aerodynamic actuators, elevators and rudders, must be considered to take into account the existing coupling effects between the three body axis. The proposed approach may be applied specially in the case in which some of the actuators are failed (stuck to a particular deflection or flapping in the wind stream) while theirs limitations, such as position, speed and response time, are considered explicitly in this study.

This recounts the story of an Iowa-born Wild West performer whose birth name was Franklin Samuel Cowdery (1867-1913). He initially gained some prominence in the US as a trick roper, marksman, and cowboy, who bore a persuasive resemblance to William F. Cody or Buffalo Bill. When he took Cody's surname, he was accused of fraudulent publicity. In 1890, he moved to Europe, where he continued to perform, but also developed an interest in observation kites and early airplane research. By the time of his early death (at age 45 in 1913), in a fatal airplane crash, he had become a public celebrity in the UK—not for his likeness to Cody, but for his aeronautic research.

Three planar, untwisted wings with the same elliptical chord distribution but with dierent curvatures of the quarter-chord line were tested in the Langley 8-Foot Transonic Pressure Tunnel (8-Ft TPT) and the Langley 7-by 10-Foot High-the tip of the crescent wing is swept outboard with a signicant separated ow region at the trailing edge of the tip. The ow at the tip of the elliptical wing with the straight trailing-edge and the ow at the tip of the crescent-shaped wing are probably inuenced by a vortex originating at the highly swept leading edges.

Three planar, untwisted wings with the same elliptical chord distribution but with dierent curvatures of the quarter-chord line were tested in the Langley 8-Foot Transonic Pressure Tunnel (8-Ft TPT) and the Langley 7-by 10-Foot High-the tip of the crescent wing is swept outboard with a signicant separated ow region at the trailing edge of the tip. The ow at the tip of the elliptical wing with the straight trailing-edge and the ow at the tip of the crescent-shaped wing are probably inuenced by a vortex originating at the highly swept leading edges.

Original scientific paper The paper is focused on the preliminary aerodynamic analyses of a new light aircraft in symmetrical flight configurations. Initial calculations have been done using a 3D vortex lattice method. Since this method is based on inviscid flow concept, the non-linear calibration diagrams for the effectiveness and circulation of control surfaces have been derived from wind tunnel test data of an existing airplane. These calibration functions have been interpolated for a new light aircraft geometry, and analyses of lift and moment coefficients with different flaps and elevator deflections have been performed. Deflected flaps and elevator generate local lift redistributions which contribute to the induced drag, even at angles of attack when the total airplane lift is zero, which has been successfully determined by the applied CFD model. Parasite drag components, omitted by CFD calculations, have been obtained using DVL and Datcom methods and superimposed with CFD results, giving complete new aircraft polars, which have been successfully applied in longitudinal stability and basic performance estimates.

Original scientific paper The paper is focused on the preliminary aerodynamic analyses of a new light aircraft in symmetrical flight configurations. Initial calculations have been done using a 3D vortex lattice method. Since this method is based on inviscid flow concept, the non-linear calibration diagrams for the effectiveness and circulation of control surfaces have been derived from wind tunnel test data of an existing airplane. These calibration functions have been interpolated for a new light aircraft geometry, and analyses of lift and moment coefficients with different flaps and elevator deflections have been performed. Deflected flaps and elevator generate local lift redistributions which contribute to the induced drag, even at angles of attack when the total airplane lift is zero, which has been successfully determined by the applied CFD model. Parasite drag components, omitted by CFD calculations, have been obtained using DVL and Datcom methods and superimposed with CFD res...

FanWing design notes 1) Introduction What follows is knowledge that was acquired over the years. In total 12 flying models have been built and three sessions were done in wind tunnels. Many design points are critical to produce an efficient lift and propulsion wing. It is not enough to put a cross flow fan in the upper side of a wing. Note also that CFD cannot simulate the operation of the wing section. This became clear in the UE funded session at the Von Karman institute in Belgium where the CFD results and the wind tunnel results were very different. 2) The rotor The rotor design is important. It is quite different from the rotors in a standard tangential/cross flow fan. The blade angle is important. After a lot of trial and error the best angle appears to be 7°. The chord of the blades is about 12% of the rotor diameter. The airfoil for the blades is hard to pin down but it must be understood that it should be optimized so that it is most efficient for the exit area of the rotor where the velocity of the air is at a maximum. The blades consequently have the trailing edge of their airfoil at the outside of the rotor. The camber of the blades is approximately same as the outer curvature of the rotor. See the diagram below. The rotor should also have a twist to reduce the blade noise. An important mechanical consideration is that the rotor should have a carbon fiber central shaft. Aluminum shafts can bend slightly on landing and are then impossible to straighten out. The rotors should have a free wheeling Sprag clutch so that if the engine stops the aircraft can autorotate.

The present paper summarizes the key findings of a numerical study on the aerodynamic behavior of a generic light aircraft under flow conditions, which are typically associated with spin motions. These conditions include high angles of attack, crosswind and rotational motion. Besides investigating the limits of Reynolds-averaged Navier-Stokes simulations for this type of conditions, the simulations contribute to a better understanding of existing wind-tunnel data. Furthermore, the influence of wind-tunnel related aspects such as scaling effects are investigated. The comparison between the numerical and experimental results generally shows good agreement regarding global force coefficients and separation behavior. The characteristic behavior with regards to crosswind and rotational motion around the longitudinal wind axis is captured well by the computations. Scaling effects are considerable with the maximum vertical force coefficient increasing by up to ∆C Z,max = 0.25 at real flight Reynolds and Mach number compared to wind-tunnel conditions.

Carburetor plays an important role in SI engine performance which supplies combustible fuel-air mixture by mixing the right amount of fuel with air before admission to the cylinder at different speed and loads. One of the important factors that affect the fuel consumption is the design of the carburetor. The carburetor provides required pressure drop with the venturi shape of the carburetor throat. Currently, alternative fuels like LPG, CNG, etc. are gaining attention all over the world because of their eco-friendly nature. So the design of the carburetor is important. Literature shows that to get a better economy and uniform distribution of fuel air mixture, it is required to design the carburetor with an effective analytical tool or software. Main objective of this work is to find pressure drop and velocity profile for different aerodynamic shape throttle valve angle and at best fuel discharge nozzle angle with help of CFD (Ansys) analysis on a carburetor and analyze for optimum design of a carburetor.

Turbulent Flow Over a Low-Camber Pitching Arc Wing MAJID MOLKI, Southern Illinois University Edwardsville -Aerodynamics of pitching airfoils and wings are of great importance to the design of air vehicles. This investigation presents the effect of camber on flow field and force coefficient for a pitching circular-arc airfoil. The wing considered in this study is a cambered plate of zero thickness which executes a linear pitch ramp, hold and return of 45 • amplitude. The momentum equation is solved on a mesh that is attached to the wing and executes a pitching motion with the wing about a pivot point located at 0.25-chord or 0.50-chord distance from the leading edge. Turbulence is modeled by the k -ω SST model. Using the open-source software OpenFOAM, the conservation equations are solved on a dynamic mesh and the flow is resolved all the way to the wall (y + ≈ 1). The computations are performed for Re = 40,000 with the reduced pitch rate equal to K = c θ/2U ∞ = 0.2. The results are presented for three wings, namely, a flat plate (zero camber) and wings of 4% and 10% camber. It is found that the flow has complex features such as leading-edge vortex, near-wake vortex pairs, clockwise and counter-clockwise vortices, and trailing-edge vortex. While vortices are formed over the flat plate, they are formed both over and under the cambered wing.

Based on the book "Convection Heat Transfer" by Adrian Bejan, this study examines the transition time of flow from laminar to turbulent in natural convection near turbulent walls. The presented analysis investigates the hydrodynamic stability of an inviscid jet near a wall. The analysis shows that if the mixing length exceeds a certain value, the transition from laminar to turbulent flow will occur. Using the scaling method, it is demonstrated that the oscillation period of the unstable jet is of the same order as the viscous penetration time perpendicular to the jet, and the transition occurs when the Reynolds number of the jet is around 100.

The shooting method is a numerical technique used to solve boundary value problems. In the Blasius equation, which is a third-order nonlinear ordinary differential equation, the shooting method can be an efficient tool for finding the velocity distribution in fluid flow over a flat plate. To use the shooting method for solving this equation, the boundary value problem must first be converted into an initial value problem. This is typically done by guessing an appropriate value for the initial derivatives at the boundary, and then using an IVP solver, such as the Runge-Kutta method, to solve the problem. By solving the Blasius equation, velocity profiles, streamlines, and thermal lines can be extracted. In all the equations, the properties of air were considered. This study is based on the book "Convection heat transfer" by Adrian Bejan. The codes for this project can be found in the appendix section of the file, written in MATLAB.

The problem described in the paper concerns the choice strategy of so-called design point of the flow engine in a multi-role aircraft at the initial stage of aircraft and engine design as the aviation system. The design point is defined by the height and speed of flight and engine parameters of heat flow which allows in particular to determine the mass and dimensions of the engine. The following analysis represents an attempt to seek other than the classic (based on a maximum within unitary thrust, specific fuel consumption) criteria for calculation point for the multi-role aircraft. Multipurpose aircraft, during every mission, very often must perform many tasks and at the same time must use the energy source for the maneuvers. The mathematical model of the chosen tasks of an aircraft has been presented, which due to the energetic requirements do not allow to build the uniform optimization criteria. The models of such flight stages have been presented: take off, climbing with the ma...

Analysis of the flow passing cylindrical obstacles is one of the basic issues in fluid dynamics and is of great importance. Many surveys have been conducted to investigate the velocity field in potential and viscous flows, as a basis for finding pressure field, and investigation of forces exerted by fluid on obstacles such as uplift and drag forces in different flow regimes. Since the nature of formation of vortexes behind the obstacles is absolutely depended on time, the conventional turbulent models which are based on average time, do not justify, and the use of direct solution of Navier Stocks equations is inevitable. One of the presented models which can solve time-dependent Navier Stocks equations in wide range of Reynolds numbers, is Random Vortex Method (RVM). Since in this method velocity field is instantly calculated, it can be used to simulate turbulent flows with a time-dependent nature. In this paper, vorticity equations gained from Navier Stocks equations are solved in both convection and diffusion phases. In this study, the flow on three cylinders with Re=140000 is investigated and field of average and instant velocity is shown along with streamlines. Also by drawing instant distribution of vortex fields and streamlines, it is possible to provide a revealing presentation of vortex behind cylinders.

Drag reduction has a significant impact on both fuel consumption and an aircraft's aerodynamic performance. The main motive of this paper is to increase the lift and decrease the drag by using shark denticles on the upper surface of the wing with different angles of attack. The shark denticle creates turbulence flow over the wing, and the boundary layer separation in the chord can be extended up to a specified distance. The NACA 0012 airfoil has been utilized to design the wing. After creating the wing, ANSYS 2022 R1 was used to CFD analyze the wing. The SST k-omega turbulence model has used the utilized RANS equation, and Mach 0.6 is assumed as free stream velocity. The findings are noteworthy, and other changes may be necessary to boost performance.

The airplane was not essentially an American invention; it was a World invention. One cannot attribute the invention of the airplane solely to the Wright brothers. The Wright brothers were voracious readers of a variety of articles written by or about other great names in aviation development, among them

Anatomy of flight vehicles, components of an airplanes and their function, configuration of space vehicle, earth's atmosphere and gravitational field, bluff bodies v/s streamlined body, airfoil. lift generation, significance of L/D ratio, aerodynamic forces, pressure, Actual and theoretical PV diagrams of four stroke and two stroke IC Engines.

The growth of aviation and the urgent need to reduce fuel consumption and emissions demand increased airspace and airport capacity and a focus on providing the preferred trajectory to each aircraft. Aircraft operators also require efficiency gains via approaches with the lowest possible minima and the significant safety benefits of vertical guidance. GNSS-based services can meet these goals and have already provided significant safety and efficiency benefits to many aircraft operators worldwide (USA, and EU-members mainly). Following this trend, Ukraine started GNSS implementation in aviation as well. Paper [1] addresses main aspects of Performance Based Navigation (PBN) implementation in Ukraine. According the requirements and methodology provided in the Manual on Testing of Radio Navigation Aids [2,3], GNSS monitoring and testing should be performed, to confirm the ability of GNSS signals to support flight procedures in accordance with Annex 10 [4]. The safe operation of an aircra...

An exposition of Bernoulli's Theorem is made, including the differential and integral approaches. Compressible and incompressible, viscous and non-viscous flows are included. In the integral approach, a stream tube is treated and global quantities are defined particularly for the viscous terms without wall slip. In duct flow, a description of flows in laminar and turbulent regimes is made.

Deep-stall is a flow regime which may occur in tandem airfoil configurations at high angle of attack. It results from a detached flow on the downstream airfoil and a drop of the lift, rendering the elevators ineffective for a potential recovery. The present numerical simulation describes vortices interaction with the downstream airfoil and its effect on change in lift for a given tandem airfoil configuration. I INTRODUCTION The phenomenon of stall is well-known for a single airfoil and leads to a collapse of the lift and a dramatic increase of the drag. On T-tail airplane configuration, where the upstream airfoil is the wing and the downstream one the tail, deep-stall occurs for high angles of attack when the tail is inside the turbulent wake of the stalled main wing. It is a stable equilibrium, where the elevators are ineffective for a potential recovery. This behavior has been described [1], [2] after the crash of the BAC 1-11 in 1963. As deep-stall is essentially two-dimensional,...

We report the effects of posture and morphology on the static aerodynamic stability and control effectiveness of physical models based on the feathered dinosaur, †Microraptor gui, from the Cretaceous of China. Postures had similar lift and drag coefficients and were broadly similar when simplified metrics of gliding were considered, but they exhibited different stability characteristics depending on the position of the legs and the presence of feathers on the legs and the tail. Both stability and the function of appendages in generating maneuvering forces and torques changed as the glide angle or angle of attack were changed. These are significant because they represent an aerial environment that may have shifted during the evolution of directed aerial descent and other aerial behaviors. Certain movements were particularly effective (symmetric movements of the wings and tail in pitch, asymmetric wing movements, some tail movements). Other appendages altered their function from creating yaws at high angle of attack to rolls at low angle of attack, or reversed their function entirely. While †M. gui lived after †Archaeopteryx and likely represents a side experiment with feathered morphology, the general patterns of stability and control effectiveness suggested from the manipulations of forelimb, hindlimb and tail morphology here may help understand the evolution of flight control aerodynamics in vertebrates. Though these results rest on a single specimen, as further fossils with different morphologies tested, the findings here could be applied in a phylogenetic context to reveal biomechanical constraints on extinct flyers arising from the need to maneuver.

Periodic and quasi-periodic time-spectral methods are applied to the demanding problems of aeroelastic flutter and gust response. Both a standard time-implicit method and a quasi-periodic time-spectral method are developed that take into account the coupling among the three fundamental aspects of computational aeroelastic calculations: unsteady flow equations, time dependent structural response to aerodynamic loads, and moving meshes. Results from the timeimplicit and time-spectral methods are then compared in order to demonstrate the capability of time-spectral methods to solve aeroelastic and gust response problems. Results from many different coupled fluid/structure problems are presented for two-dimensional representations of both fixed-wing and rotary-wing aeroelastic response both with and without gusts of different lengths. It is shown that time-spectral based methods can be used to solve every reasonable combination of airfoil motion and gust response that has thus far been ...

Wind power is the source of propulsive energy for sailboats and albatrosses. We present the UNAv, an Unmanned Nautical Air-water vehicle, that borrows features from both. It is composed of a glider-type airframe fitted with a vertical wing-sail extending above the center of mass of the system and a vertical surface-piercing hydrofoil keel extending below. The sail and keel are both actuated in pitch about their span-wise axes. Like an albatross, the UNAv is fully streamlined, high lift-to-drag ratio and generates the gravity-cancelling force by means of its airborne wings. Like a sailboat, the UNAv interacts with water and may access the full magnitude of the wind. A trim analysis predicts that a 3.4-meter span, 3 kg system could stay airborne in winds as low as 2.8 m/s (5.5 knots), and travel several times faster than the wind speed. Trim flight requires the ability to fly at extreme low height with the keel immersed in water. For that purpose, a multi-input longitudinal flight controller that leverages fast flap actuation is presented. The flight maneuver is demonstrated experimentally.

In order to meet the goals of reducing noise, emissions, and fuel burn, the Hybrid Wing Body (HWB) will require revolutionary improvements in aerodynamics, acoustics, and structural design. Reducing or eliminating the traditional "tube" design fuselage offers a significant improvement in aerodynamic efficiency. The use of embedded engines offers the possibility of reducing radiated noise to the community while at the same time alleviating separation of the airflow over the wing. Reducing or eliminating empennage surfaces offers the possibility to reduce skin friction, form and trim drag during cruise while reducing structural weight. Essentially, the HWB offers the potential for significant improvements by using a multidisciplinary , highly integrated approach to aircraft design. For In designing HWB concepts that includes Cruise Efficient Short TakeOff and Landing (CESTOL) characteristics, some performance issues have been identified. The tendency for the powered lift system to generate nose down pitching moments makes it difficult to trim the HWB at low takeoff speeds without a horizontal tail, however the inclusion of a tail reduces efficiency due to the short longitudinal moment arm a HWB aircraft typically would have. Another issue stems from the desire for the HWB aircraft to use buried engines for noise mitigation purposes, but the need to efficiently vector thrust for short field performance means that a nacelle or off body mounted propulsion system may be preferred. Not all of the design integration issues are detrimental. For example, the large internal wing volume of the HWB provides more space for the powered-lift system than conventional tube-with-wing configurations with smaller wing areas. This paper will examine some of these issues of integrating CESTOL performance into a HWB aircraft. It will look at the potential advantages of a HWB CESTOL, examine areas of conflicting requirements, and offer potential solutions to mitigating these conflicts, or at least present a possible path to a solution. It will also look at the process being used in NASA's Subsonic Fixed Wing Project to develop new tools and modeling techniques for developing new designs for HWB CESTOL and validating those tools through experiments and wind tunnel tests in cooperation with industry partners and through NASA Research Announcements (NRA).

In order to meet the goals of reducing noise, emissions, and fuel burn, the Hybrid Wing Body (HWB) will require revolutionary improvements in aerodynamics, acoustics, and structural design. Reducing or eliminating the traditional "tube" design fuselage offers a significant improvement in aerodynamic efficiency. The use of embedded engines offers the possibility of reducing radiated noise to the community while at the same time alleviating separation of the airflow over the wing. Reducing or eliminating empennage surfaces offers the possibility to reduce skin friction, form and trim drag during cruise while reducing structural weight. Essentially, the HWB offers the potential for significant improvements by using a multidisciplinary , highly integrated approach to aircraft design. For In designing HWB concepts that includes Cruise Efficient Short TakeOff and Landing (CESTOL) characteristics, some performance issues have been identified. The tendency for the powered lift system to generate nose down pitching moments makes it difficult to trim the HWB at low takeoff speeds without a horizontal tail, however the inclusion of a tail reduces efficiency due to the short longitudinal moment arm a HWB aircraft typically would have. Another issue stems from the desire for the HWB aircraft to use buried engines for noise mitigation purposes, but the need to efficiently vector thrust for short field performance means that a nacelle or off body mounted propulsion system may be preferred. Not all of the design integration issues are detrimental. For example, the large internal wing volume of the HWB provides more space for the powered-lift system than conventional tube-with-wing configurations with smaller wing areas. This paper will examine some of these issues of integrating CESTOL performance into a HWB aircraft. It will look at the potential advantages of a HWB CESTOL, examine areas of conflicting requirements, and offer potential solutions to mitigating these conflicts, or at least present a possible path to a solution. It will also look at the process being used in NASA's Subsonic Fixed Wing Project to develop new tools and modeling techniques for developing new designs for HWB CESTOL and validating those tools through experiments and wind tunnel tests in cooperation with industry partners and through NASA Research Announcements (NRA).

The aerodynamic performance of a wing at an oblique deployment orientation has been found through wind tunnel testing to affect both the lateral and longitudinal stability of a cruise missile. While conventional analysis tools are insufficient for calculating the aerodynamics of an oblique wing, a suitable method has been determined for use with the USAF Missile DATCOM code. Comparisons made between wind tunnel results and Missile DATCOM calculations show that Missile DATCOM can produce a reasonable approximation of the aerodynamics of a wing at oblique deployment angles.

Current aviation standards define a multipath error model that is valid after the smoothing filter is assumed to have converged (assuming a 100 s Hatch filter). The draft standards for dual frequency Satellite-based Augmentation Systems further specify an error model when the code has not been smoothed, and it is defined as a multiple of the converged value. In this paper, multipath and noise error bounds are derived as a function of smoothing time assuming a first order model for the code multipath and the receiver noise. These error bounds are evaluated using GPS and Galileo measurements collected in flight. The derived model appears to account well for the error reduction as a function of smoothing time.

graduated as an electronics engineer in 1992 from ENAC (Ecole Nationale de l'Aviation Civile) in Toulouse, France. Since 1994, he has been working on the application of satellite navigation techniques to civil aviation. He received his Ph.D. in 1997 and has been in charge of the signal processing lab of the ENAC since 2000. Igor V. Nikiforov received the M.S. degree in automatic control from the Moscow Physical-Technical Institute in 1974, and the Ph.D. in automatic control from the Institute of Control Sciences, Moscow, in 1981. He joined the University of Technology of Troyes in 1995, France, where he is professor and head of the Institute of Computer Sciences and Engineering of Troyes, FRE CNRS 2732. His scientific interests include statistical decision theory, fault detection, signal processing, and navigation. Lionel Fillatre received the Ph.D. degree in systems optimization from the UTT-the University of Technology of Troyes, France, in 2004. He is currently an assistant professor at the Institute of Computer Sciences and Engineering of Troyes (ISTIT), FRE CNRS 2732. His current research interests include decision theory, signal and image processing and non-destructive testing. Benoit Roturier graduated as a CNS systems engineer from Ecole Nationale de l'Aviation Civile (ENAC), Toulouse in 1985 and obtained a PhD in Electronics from Institut National Polytechnique de Toulouse in 1995. He was successively in charge of Instrument Landing Systems at DGAC/STNA (Service Technique de la Navigation Aerienne), then of research activities on CNS systems at ENAC. He is now head of GNSS Navigation subdivision at STNA and is involved in the development of civil aviation applications based on GPS/ABAS, EGNOS and Galileo. Eric Chatre graduated as an electronics engineer in 1992 from the ENAC (Ecole Nationale de l'Aviation Civile), Toulouse, France. From 1994 to 2001, he worked with the Service Technique de la Navigation Aérienne (STNA) in Toulouse on implementation of satellite navigation in civil aviation. He is now part of the Galileo Joint Undertaking where he is in charge of standardisation and certification matters. He is secretary of Eurocae WG 62. Mathieu Raimondi will graduate in 2005 as an electronics engineer from the Ecole Nationale de l'Aviation Civile (ENAC) in Toulouse, France. Anne-Christine Escher graduated as an electronics engineer in 1999 from ENAC (Ecole Nationale de l'Aviation Civile) in Toulouse, France. Since 1999, she has been working on the application of satellite navigation techniques to civil aviation at the signal processing lab of the ENAC.

T HIS Synoptic describes an analytical method for solving the equation governing the inviscid, irrotational, compressible, potential flow about a propeller. The equation and boundary conditions are transferred to a noninertial system of coordinates rotating with the propeller, in which the basic problem becomes a steady one. The solution method takes advantage of the linearity of the model by superposing a "compressible" solution to the potential equation on an "incompressible" wake solution. In addition, the boundary conditions are satisfied by dividing the flowfield at the propeller plane, solving the equations separately ahead of and behind this plane, and enforcing continuity matching conditions. Applying the final boundary condition yields an infinite-series integral equation for the unknown circulation distribution. A lifting-line method is used to produce numerical results. Presented results establish the effect of compressibility on the induced field. Contents Except in the vicinity of the propeller and its wake, the flowfield caused by an advancing propeller in an inertial coordinate system may be considered irrotational and isentropic. This approach treats the viscous drag on the blade as inconsequential to the general induction theory. Consider a transformation of the linearized convective wave equation to a bladefixed coordinate system that rotates with the propeller at angular velocity w. This noninertial, cylindrical system (r, 6, z, t) is defined by r = r, 6 = 0 + ui, z = z, and t = t. Tildes represent quantities in the inertial, translating coordinate system. Performing these coordinate transformations gives the unsteady, linearized potential equation in the rotating system. Although the general methodology works for unsteady flows, this Synoptic focuses on the aerodynamics of the steady propeller, that is, a rigid propeller experiencing a uniform freestream. This allows the time-dependent terms to be excluded, since the flow now appears steady in the rotating, translating coordinates. Converting the dimensional spatial variables to dimensionless form (p = a>/-/£/and z = uz/U) and dropping the bar over the dimensionless z variable, the steady, potential equation becomes-M 2 p 2 (1) with /3 2 = 1-M 2 and M = (u/a) is the advancing Mach number. This linear analysis assumes that all motion of the helicoidal wake occurs in the axial direction and that no distortion of the

has over 35 years of aerospace, utility, commercial, and industrial experience as an engineer, technical manager, program manager, corporate executive, and consultant. He received the Society of Automotive Engineer's Outstanding Presentation Award and the Honeywell Business Growth Award. In the past he has taught engineering and business administration courses at several universities. He is currently an associate faculty member of the

The Pacific Basin, by virtue of its vastness and its complex aeroscape, provides unique opportunities to address questions about the behavioral and physiological capabilities and mechanisms through which birds can complete spectacular flights. No longer is the Pacific seen just as a formidable barrier between terrestrial habitats in the north and the south, but rather as a gateway for specialized species, such as shorebirds, to make a living on hemispherically distributed seasonal resources. This recent change in perspective is dramatic, and the research that underpins it has presented new opportunities to learn about phenomena that often challenge a sense of normal. Ancient Polynesians were aware of the seasonal passage of shorebirds and other landbirds over the Pacific Ocean, incorporating these observations into their navigational “tool kit” as they explored and colonized the Pacific. Some ten centuries later, systematic visual observations and tracking technology have revealed m...

In this paper we discuss the first results from a set of flight experiments. The objective was to show that terrestrial ranging can support aviation services with sufficient accuracy for non-precision approach. The experimental setup is based on using LDACS signals to provide pseudorange measurements. The data analysis reveals that the position solutions are compliant with RNP 1 standards. We also propose the feasibility of RNP 0.3 using LDACS as terrestrial ranging.

This paper proposes one possible concept for a dual frequency dual constellation GBAS architecture. It is based on a single frequency L5/E5a mode as primary processing scheme for best standard performance, a switch to an ionosphere free combination in case of ionospheric disturbances and supporting also classical GBAS approach service types (GAST) C and D for single frequency GPS-based CAT I and CAT II/III modes. The concept is supported by a proposal of how to transmit the required corrections in the existing capacity limited VDB broadcast and is backwards compatible to legacy GBAS. A discussion about the benefits and remaining issues of the proposed architecture concludes the paper.

The aerodynamic performance of a wing at an oblique deployment orientation has been found through wind tunnel testing to affect both the lateral and longitudinal stability of a cruise missile. While conventional analysis tools are insufficient for calculating the aerodynamics of an oblique wing, a suitable method has been determined for use with the USAF Missile DATCOM code. Comparisons made between wind tunnel results and Missile DATCOM calculations show that Missile DATCOM can produce a reasonable approximation of the aerodynamics of a wing at oblique deployment angles.

Kapal mengalami keterlambatan dalam proses pengangkutan muatan semen curah sehingga mengalami waktu tunggu yang dapat mempengaruhi target pencapaian pengangkutan muatan. Penelitian ini bertujuan untuk mengetahui pengaruh ketersediaan muatan kapal dan waiting time terhadap volume muatan kapal yang diangkut di kapal milik PT. Indobaruna Bulk Transport. Tidak tercapainya target pengangkutan volume muatan kapal per tahun dikarenakan keterlambatan ketersediaan muatan kapal dan angka waiting time yang tinggi sehingga target pengangkutan muatan tidak tercapai. Penelitian ini dilakukan selama 3 bulan dengan meneliti sebanyak 7 kapal dengan jenis voyage charter milik PT. Indobaruna Bulk Transport yang seluruhnya merupakan kapal bermuatan semen curah. Hasil penelitian yang diperoleh dengan α = 0,05 adalah: (1) adanya hubungan yang positif dan sangat kuat antara ketesediaan muatan kapal terhadap volume muatan kapal yang diangkut di kapal milik PT. Indobaruna Bulk Tansport, dimana rx1y =0,92 dengan thitung=5,125 dan koefisien determinasi sebesar 84%, (2) adanya hubungan yang positif dan kuat antara waiting time terhadap volume muatan kapal yang diangkut di kapal milik PT. Indobaruna Bulk Transport, dimana rx2y=0,76 dengan thitung=3,930 serta koefisien determinasi sebesar 57%, (3) adanya hubungan yang kuat antara ketersediaan muatan kapal dan waiting time terhadap volume muatan kapal yang diangkut di kapal milik PT. Indobaruna Bulk Transport, dimana rx1x2y=0,94 dengan Fhitung=7,25 serta koefisien determinasi sebesar 88%.

Although quasi-steady aerodynamic modelling is a widely used technique, its shortcomings in representing the stall and post-stall dynamics have been noted in the literature. Various methods to model the unsteady aerodynamics effects have been proposed as a result, but their direct implications on the aircraft's flight dynamics characteristics have not been widely studied. In this paper, we combine the state-space method for unsteady aerodynamic modelling with bifurcation analysis to examine the sensitivity of stall and post-stall behaviour to the choice of aerodynamic modelling method: quasi-steady or unsteady. It is found that quasi-steady modelling is adequate for the chosen example of a T-tailed transport aircraft that does not undergo rapid manoeuvring. The study is then expanded to investigate a hypothetical situation with highly unsteady aerodynamic characteristics resembling a delta wing configurationachieved by increasing the time delay constants in the unsteady model. This results in an aircraft with significantly lower flying qualities as indicated by bifurcation analysis. These findings highlight the need to implement unsteady aerodynamic modelling techniques in high-performance aircraft with significant vortex-related unsteady aerodynamics in order to sufficiently capture their stall and post-stall dynamics.

Although there have been many proposed methods to model unsteady aerodynamic effects in the stall and post-stall region, little work has been done to directly assess the impact of unsteady aerodynamic models on stability and control characteristics. In this paper, we combine the state-space method for unsteady aerodynamic modelling with bifurcation analysis to examine the sensitivity of stall and post-stall behaviour to the choice of aerodynamic modelling method: quasi-steady or unsteady. It is found that quasi-steady modelling can adequately capture the dynamics of the chosen example of a T-tailed transport aircraft with negligible wing-tail coupling. The study is then expanded to investigate a hypothetical situation with highly unsteady aerodynamic characteristics resembling a delta wing configurationachieved by increasing the time delay constants in the unsteady model. This results in an aircraft with significantly lower flying qualities as indicated by bifurcation analysis. These findings highlight the need to implement unsteady aerodynamic modelling techniques in high-performance aircraft with significant vortex-related unsteady aerodynamics in order to sufficiently capture their stall and post-stall dynamics.

Management 86-87 XIV. Maintainability 87-89 XV. Design Schedules 90-91 XVI. Cost 92 XVII. RFP Requirements 93-94 XVIII. Conclusion 95 6-Sideslip angle ABSTRACT At the equator, the ozone layer ranges from 65,000 to 130,000+ feet which is beyond the capabilities of the ER-2, NASA's current high altitude reconnaissance aircraft. The Universities Space Research Association, in cooperation with NASA, is sponsoring an undergraduate program which is geared to designing an aircraft that can study the ozone layer at the equator. This aircraft must be able to satisfy four mission profiles. Mission one is a polar mission which ranges from Chile to the South Pole and back to Chile, a total range of 6000 n. mi at 100,000 feet with a 2500 Ib. payload. The second mission is also a polar mission with a decreased altitude of 70,000 feet and an increased payload of 4000 Ib.

Management 86-87 XIV. Maintainability 87-89 XV. Design Schedules 90-91 XVI. Cost 92 XVII. RFP Requirements 93-94 XVIII. Conclusion 95 6-Sideslip angle ABSTRACT At the equator, the ozone layer ranges from 65,000 to 130,000+ feet which is beyond the capabilities of the ER-2, NASA's current high altitude reconnaissance aircraft. The Universities Space Research Association, in cooperation with NASA, is sponsoring an undergraduate program which is geared to designing an aircraft that can study the ozone layer at the equator. This aircraft must be able to satisfy four mission profiles. Mission one is a polar mission which ranges from Chile to the South Pole and back to Chile, a total range of 6000 n. mi at 100,000 feet with a 2500 Ib. payload. The second mission is also a polar mission with a decreased altitude of 70,000 feet and an increased payload of 4000 Ib.