Dissertations / Theses: 'Attack angle'

1

Sims, J. P. "In-flight measurement of angle of attack." Thesis, Swansea University, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.639041.

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Three possible techniques for measuring angle of attack (AOA) have been investigated with a view to establishing a low-cost method for use in light aircraft. Basically, methods using new silicon bonded pressure transducers and low cost inertial components have been compared with a standard AOA system. An Inertial Measurement Unit (IMU) and a differential pressure AOA system have been designed and constructed using the low-cost transducers and both have been installed in a twin engined general aviation aircraft along with the Teledyne AOA cone probe. A series of flight tests have been conducted to analyze and compare the three AOA sensors and the results have shown that the low-cost differential pressure sensor system provided data as good as the more expensive commercial unit for approximately one-tenth of its cost. The IMU system required high accuracy in the acceleration channels and though this precluded the use of low-cost accelerometers, the low-cost rate and attitude sensors coupled with the high accuracy accelerometers did provide adequate AOA data. The final results demonstrated that the differential pressure AOA was accurate to within ±0.50° and the low-cost IMU (with high accuracy acceleration calculated AOA) was accurate to within ±1.50°.

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2

Hoang, Ngoc T. "The hemisphere-cylinder at an angle of attack." Diss., This resource online, 1991. http://scholar.lib.vt.edu/theses/available/etd-08062007-094404/.

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3

Atesoglu, Ozgur Mustafa. "High Angle Of Attack Maneuvering And Stabilization Control Of Aircraft." Phd thesis, METU, 2007. http://etd.lib.metu.edu.tr/upload/12608575/index.pdf.

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In this study, the implementation of modern control techniques, that can be used both for the stable recovery of the aircraft from the undesired high angle of attack flight state (stall) and the agile maneuvering of the aircraft in various air combat or defense missions, are performed. In order to accomplish this task, the thrust vectoring control (TVC) actuation is blended with the conventional aerodynamic controls. The controller design is based on the nonlinear dynamic inversion (NDI) control methodologies and the stability and robustness analyses are done by using robust performance (RP) analysis techniques. The control architecture is designed to serve both for the recovery from the undesired stall condition (the stabilization controller) and to perform desired agile maneuvering (the attitude controller). The detailed modeling of the aircraft dynamics, aerodynamics, engines and thrust vectoring paddles, as well as the flight environment of the aircraft and the on-board sensors is performed. Within the control loop the human pilot model is included and the design of a fly-by-wire controller is also investigated. The performance of the designed stabilization and attitude controllers are simulated using the custom built 6 DoF aircraft flight simulation tool. As for the stabilization controller, a forced deep-stall flight condition is generated and the aircraft is recovered to stable and pilot controllable flight regimes from that undesired flight state. The performance of the attitude controller is investigated under various high angle of attack agile maneuvering conditions. Finally, the performances of the proposed controller schemes are discussed and the conclusions are made.

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4

Ganji, Farid 1967. "Control of the space shuttle angle-of-attack during reentry." Thesis, Massachusetts Institute of Technology, 2002. http://hdl.handle.net/1721.1/89340.

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5

Walter, Daniel James, and Daniel james walter@gmail com. "Study of aerofoils at high angle of attack in ground effect." RMIT University. Aerospace, Mechanical and Manufacturing Engineering, 2007. http://adt.lib.rmit.edu.au/adt/public/adt-VIT20080110.145138.

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Aerodynamic devices, such as wings, are used in higher levels of motorsport (Formula-1 etc.) to increase the contact force between the road and tyres (i.e. to generate downforce). This in turn increases the performance envelope of the race car. However the extra downforce increases aerodynamic drag which (apart from when braking) is generally detrimental to lap-times. The drag acts to slow the vehicle, and hinders the effect of available drive power and reduces fuel economy. Wings, in automotive use, are not constrained by the same parameters as aircraft, and thus higher angles of attack can be safely reached, although at a higher cost in drag. Variable geometry aerodynamic devices have been used in many forms of motorsport in the past offering the ability to change the relative values of downforce and drag. These have invariably been banned, generally due to safety reasons. The use of active aerodynamics is currently legal in both Formula SAE (engineering compet ition for university students to design, build and race an open-wheel race car) and production vehicles. A number of passenger car companies are beginning to incorporate active aerodynamic devices in their designs. In this research the effect of ground proximity on the lift, drag and moment coefficients of inverted, two-dimensional aerofoils was investigated. The purpose of the study was to examine the effect ground proximity on aerofoils post stall, in an effort to evaluate the use of active aerodynamics to increase the performance of a race car. The aerofoils were tested at angles of attack ranging from 0° - 135°. The tests were performed at a Reynolds number of 2.16 x 105 based on chord length. Forces were calculated via the use of pressure taps along the centreline of the aerofoils. The RMIT Industrial Wind Tunnel (IWT) was used for the testing. Normally 3m wide and 2m high, an extra contraction was installed and the section was reduced to form a width of 295mm. The wing was mounted between walls to simulate 2-D flow. The IWT was chosen as it would allow enough height to reduce blockage effect caused by the aerofoils when at high angles of incidence. The walls of the tunnel were pressure tapped to allow monitoring of the pressure gradient along the tunnel. The results show a delay in the stall of the aerofoils tested with reduced ground clearance. Two of the aerofoils tested showed a decrease in Cl with decreasing ground clearance; the third showed an increase. The Cd of the aerofoils post-stall decreased with reduced ground clearance. Decreasing ground clearance was found to reduce pitch moment variation of the aerofoils with varied angle of attack. The results were used in a simulation of a typical Formula SAE race car.

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6

Petterson, Kristian. "The aerodynamics of slender aircraft forebodies at high angle of attack." Thesis, Cranfield University, 2001. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.392234.

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7

Cohen, David E. II. "Trim Angle of Attack of Flexible Wings Using Non-Linear Aerodynamics." Diss., Virginia Tech, 1998. http://hdl.handle.net/10919/30404.

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Multidisciplinary interactions are expected to play a significant role in the design of future high-performance aircraft (Blended-Wing Body, Truss-Braced wing, High Speed Civil transport, High-Altitude Long Endurance aircraft and future military aircraft). Also, the availability of supercomputers has made it now possible to employ high-fidelity models (Computational Fluid Dynamics for fluids and detailed finite element models for structures) at the preliminary design stage. A necessary step at that stage is to calculate the wing angle-of-attack at which the wing will generate the desired lift for the specific flight maneuver. Determination of this angle, a simple affair when the wing is rigid and the flow regime linear, becomes difficult when the wing is flexible and the flow regime non-linear. To solve this inherently nonlinear problem, a Newton's method type algorithm is developed to simultaneously calculate the deflection and the angle of attack. The present algorithm requires the sensitivity of the aerodynamic pressure with respect to each of the generalized displacement coordinates needed to represent the structural displacement. This sensitivity data is easy to determine analytically when the flow regime is linear. The present algorithm uses a finite difference method to obtain these sensitivities and thus requires only the pressure data and the surface geometry from the aerodynamic model. This makes it ideally suited for nonlinear aerodynamics for which it is difficult to obtain the sensitivity analytically. The present algorithm requires the CFD code to be run for each of the generalized coordinates. Therefore, to reduce the number of generalized coordinates considerably, we employ the modal superposition approach to represent the structural displacements. Results available for the Aeroelastic Research Wing (ARW) are used to evaluate the performance of the modal superposition approach. Calculations are made at a fixed angle of attack and the results are compared to both the experimental results obtained at NASA Langley Research Center, and computational results obtained by the researchers at NASA Ames Research Center. Two CFD codes are used to demonstrate the modular nature of this research. Similarly, two separate Finite Element codes are used to generate the structural data, demonstrating that the algorithm is not dependent on using specific codes. The developed algorithm is tested for a wing, used for in-house aeroelasticity research at Boeing (previously McDonnell Douglas) Long Beach. The trim angle of attack is calculated for a range of desired lift values. In addition to the Newton's method algorithm, a non derivative method (NDM) based on fixed point iteration, typical of fixed angle of attack calculations in aeroelasticity, is employed. The NDM, which has been extended to be able to calculate trim angle of attack, is used for one of the cases. The Newton's method calculation converges in fewer iterations, but requires more CPU time than the NDM method. The NDM, however, results in a slightly different value of the trim angle of attack. It should be noted that NDM will converge in a larger number of iterations as the dynamic pressure increases. For one value of the desired lift, both viscous and inviscid results were generated. The use of the inviscid flow model while not resulting in a markedly different value for the trim angle of attack, does result in a noticeable difference both in the wing deflection and the span loading when compared to the viscous results. A crude (coarse-grain) parallel methodology was used in some of the calculations in this research. Although the codes were not parallelized, the use of modal superposition made it possible to compute the sensitivity terms on different processors of an IBM SP/2. This resulted in a decrease in wall clock time for these calculations. However, even with the parallel methodology, the CPU times involved may be prohibitive (approximately 5 days per Newton iteration) to any practical application of this method for wing analysis and design. Future work must concentrate on reducing these CPU times. Two possibilities: (i) The use of alternative basis vectors to further reduce the number of basis vectors used to represent the structural displacement, and (ii) The use of more efficient methods for obtaining the flow field sensitivities. The former will reduce the number of CFD analyses required the latter the CPU time per CFD analysis. NOTE: (03/2007) An updated copy of this ETD was added after there were patron reports of problems with the file.
Ph. D.

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8

Mohmad, Rouyan Nurhana. "Model simulation suitable for an aircraft at high angle of attack." Thesis, Cranfield University, 2016. http://dspace.lib.cranfield.ac.uk/handle/1826/9722.

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Simulation of a dynamic system is known to be sensitive to various factors and one of them could be the precision of model parameters. While the sensitivity of flight dynamic simulation to small changes in aerodynamic coefficients is typically not studied, the simulation of aircraft required to operate in nonlinear flight regimes usually at high angles of attack can be very sensitive to such small differences. Determining the significance and impact of the differences in aerodynamic characteristics is critical for understanding the flight dynamics and designing suitable flight control laws. This thesis uses this concept to study the effect of the differences in aerodynamic data for different aerodynamic models provided for a same aircraft which is F-18 HARV combat aircraft. The aircraft was used as a prototype for the high angles of attack technology program. However modeling an aircraft at high angles of attack requires an extensive aerodynamic data which are usually di cult to access. All aerodynamic models were collected from open literature and implemented within a nonlinear six degree of freedom aircraft model. Inspection of aerodynamic data set for these models has shown mismatches for certain aerodynamic derivatives, especially at higher angles of attack where nonlinear dynamics are known to exist. Nonlinear simulations are used to analyse three different types of flight dynamic models that use look-up-tables, arc-tangent formulation and polynomial functions to represent aerodynamic data that are suitable for high angles of attack application. To achieve this, a nonlinear six degree of freedom Simulink model was developed to accommodate these aerodynamic models separately. The trim conditions were obtained for different combinations of angles of attack and airspeed and the models were linearized in each case. Properties of the resulting state matrices such as eigenvalues and eigenvectors were studied to determine the dynamic behaviour of the aircraft at various flight conditions.

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9

Ravi, R. "High Angle of Attack Forebody Flow Physics and Design Emphasizing Directional Stability." Diss., This resource online, 1997. http://scholar.lib.vt.edu/theses/available/etd-01252008-163458/.

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10

Yip, Pui-Chuen Patrick. "A comparison of control design options for high angle-of-attack flights." Thesis, Massachusetts Institute of Technology, 1991. http://hdl.handle.net/1721.1/13431.

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Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1991, and Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Electrical Engineering and Computer Science.
Includes bibliographical references (p. 168-170).
by Pui-Chuen Patrik Yip.
M.S.

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11

Stagg, Gregory A. "An Aerodynamic Model for Use in the High Angle of Attack Regime." Thesis, Virginia Tech, 1998. http://hdl.handle.net/10919/35596.

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Harmonic oscillatory tests for a fighter aircraft using the Dynamic Plunge--Pitch--Roll model mount at Virginia Tech Stability Wind Tunnel are described. Corresponding data reduction methods are developed on the basis of multirate digital signal processing. Since the model is sting mounted, the frequencies associated with sting vibration are included in balance readings thus a linear filter must be used to extract out the aerodynamic responses. To achieve this, a Finite Impulse Response (FIR) is designed using the Remez exchange algorithm. Based on the reduced data, a state--space model is developed to describe the unsteady aerodynamic characteristics of the aircraft during roll oscillations. For this model, we chose to separate the aircraft into panels and model the local forces and moments. Included in this technique is the introduction of a new state variable, a separation state variable which characterizes the separation for each panel. This new variable is governed by a first order differential equation. Taylor series expansions in terms of the input variables were performed to obtain the aerodynamic coefficients of the model. These derivatives, a form of the stability derivative approach, are not constant but rather quadratic functions of the new state variable. Finally, the concept of the model was expanded to allow for the addition of longitudinal motions. Thus, pitching moments will be identified at the same time as rolling moments. The results show that the goal of modeling coupled longitudinal and lateral--directional characteristics at the same time using the same inputs is feasible.
Master of Science

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12

Ko, Joon Soo. "Analysis of the dynamic stability derivatives for high angle of attack aircraft." Diss., Virginia Polytechnic Institute and State University, 1985. http://hdl.handle.net/10919/52300.

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Modern, high performance aircraft are required to be able to fly and be controlled over a wide variety of flight conditions. In order to predict the aircraft behavior and control requirements over the entire flight regime it is necessary to have a proper aerodynamic model. Flight conditions at high angles of attack lead to separated flows making the aerodynamic model more difficult to obtain. In this research wind tunnel experiments are performed on an F-5 air-craft model at high angles of attack, with small oscillations about the body oriented roll axis. In addition the free stream environment can be configured in one of three ways: l) straight uniform flow, 2) curved flow to simulated a horizontal turn, and 3) rolling flow to simulated a roll motion about the relative Velocity vector.
Ph. D.

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13

Stucke, Russell Andrew. "High Angle-of-Attack Yaw Control Using Strakes on Blunt-Nose Bodies." University of Toledo / OhioLINK, 2006. http://rave.ohiolink.edu/etdc/view?acc_num=toledo1167777201.

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14

Jouannet, Christopher. "Model based aircraft design : high angle of attack aerodynamics and weight estimation methods /." Linköping : Dept. of Mechanical Engineering, Linköping University, 2005. http://www.bibl.liu.se/liupubl/disp/disp2005/tek968s.pdf.

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15

Johnson, Dan A. "Flowfield measurements in the wake of a missile at high angle of attack." Thesis, Monterey, California. Naval Postgraduate School, 1989. http://hdl.handle.net/10945/27059.

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16

Fan, Yigang. "Identification of an Unsteady Aerodynamic Model up to High Angle of Attack Regime." Diss., Virginia Tech, 1997. http://hdl.handle.net/10919/29830.

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The harmonic oscillatory tests for a fighter aircraft configuration using the Dynamic Plunge-Pitch-Roll (DyPPiR) model mount at Virginia Tech Stability Wind Tunnel are described and analyzed. The corresponding data reduction methods are developed on the basis of multirate digital signal processing techniques. Since the model is sting-mounted to the support system of DyPPiR, the Discrete Fourier Transform (DFT) is first used to identify the frequencies of the elastic modes of sting. Then the sampling rate conversion systems are built up in digital domain to resample the data at a lower rate without introducing distortions to the signals of interest. Finally linear-phase Finite Impulse Response (FIR) filters are designed by Remez exchange algorithm to extract the aerodynamic characteristics responses to the programmed motions from the resampled measurements. These data reduction procedures are also illustrated through examples. The results obtained from the harmonic oscillatory tests are then illustrated and the associated flow mechanisms are discussed. Since no significant hysteresis loops are observed for the lift and the drag coefficients for the current angle of attack range and the tested reduced frequencies, the dynamic lags of separated and vortex flow effects are small in the current oscillatory tests. However, large hysteresis loops are observed for pitch moment coefficient in the current tests. This observation suggests that at current flow conditions, pitch moment has large pitch rate and alpha-dot dependencies. Then the nondimensional maximum pitch rate q__max is introduced to characterize these harmonic oscillatory motions. It is found that at current flow conditions, all the hysteresis loops of pitch moment coefficient with same nondimensional maximum pitch rate are tangential to one another at both top and bottom of the loops, implying approximately same maximum offset of these loops from static values. Several cases are also illustrated. Based on the results obtained and those from references, a state-space model is developed to describe the unsteady aerodynamic characteristics up to the high angle of attack regime. A nondimensional coordinate is introduced as the state variable describing the flow separation or vortex burst. First-order differential equation is used to govern the dynamics of flow separation or vortex bursting through this state variable. To be valid for general configurations, Taylor series expansions in terms of the input variables are used in the determination of aerodynamic characteristics, resembling the current approach of the stability derivatives. However, these derivatives are longer constant. They are dependent on the state variable of flow separation or vortex burst. In this way, the changes in stability derivatives with the angle of attack are included dynamically. The performance of the model is then validated by the wind-tunnel measurements of an NACA 0015 airfoil, a 70 degree delta wing and, finally two F-18 aircraft configurations. The results obtained show that within the framework of the proposed model, it is possible to obtain good agreement with different unsteady wind tunnel data in high angle-of-attack regime.
Ph. D.

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17

MAY, CAMERON. "HIGH ANGLE OF ATTACK FLIGHT CONTROL OF DELTA WING AIRCRAFT USING VORTEX ACTUATORS." University of Cincinnati / OhioLINK, 2005. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1109166873.

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18

Sirangu, Vijaya. "AERODYNAMIC CONTROL OF SLENDER BODIES AT HIGH ANGLES OF ATTACK." University of Toledo / OhioLINK, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=toledo1271365316.

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19

Rosti, Marco. "Direct numerical simulation of an aerofoil at high angle of attack and its control." Thesis, City, University of London, 2016. http://openaccess.city.ac.uk/15843/.

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Detailed analysis of the flow around a NACA0020 aerofoil at moderate low chord Reynolds number (Rec = 2×104) in completely stalled conditions has been carried out by means of Direct Numerical Simulations. The stalled condition is either a steady configuration at a fixed angle of attack (α = 20o) or it is reached via a ramp-up manoeuvre, increasing the angle of attack from 0o to 20o. Concerning this last case, new insights on the vorticity dynamics leading to the lift overshoot, lift crisis and the damped oscillatory cycle that gradually matches the steady condition, are discussed using a number of post-processing techniques. These include a detailed analysis of the flow ensemble average statistics and coherent structures identification that has been carried out using the Q-criterion and the Finite-Time Lyapunov Exponent technique. Based on the fundamental knowledge achieved in studying the static and the dynamic stall, we introduced a biomimetic passive control technique to mitigate the aerodynamic performance degradation typical of such flow conditions. In particular, the envisaged control technique has been inspired by the dorsal feathers that are used by almost all birds to adapt their wing characteristics to delay stall or to moderate its adverse effects (e.g., during landing or sudden increase in angle of attack due to gusts). Some of the feathers are believed to pop up as a consequence of flow separation and to interact with the flow producing beneficial modifications of the unsteady vorticity field. The adoption of self adaptive flaplets in aircrafts, inspired by birds feathers, requires the understanding of the physical mechanisms leading to their aerodynamic benefits and the determination of the characteristics of optimal flaps including their size, positioning and ideal fabrication material. In this framework, we have used numerical simulation to study the effects of this passive control technique in both steady and dynamic stall. In particular, for the static case, we have defined an optimal condition as the one that delivers the highest lift coefficient CL, preserving or improving the aerodynamic efficiency E = CL/CD. To achieve a condition close to optimality we started by considering a simplified scenario, to determine the main characteristics of the flap (i.e., variations of its length, position and natural frequency). Later on, a detailed direct numerical simulation analysis is used to understand the origin of the aerodynamic benefits introduced by the pop-up of the optimal flaplet. It is found that an optimal flap can deliver a mean lift increase of about 20% on a NACA0020 aerofoil at an incidence of 20o degrees. The analysis of direct numerical simulation data of the flow field around the aerofoil equipped with the optimal flap allowed to elucidate the main mechanism that promotes the aerodynamic improvements. In particular, it is found that the flaplet movement, induced by the transit of a large recirculation bubble on the aerofoil suction side, displaces the trailing edge vortices further downstream, away from the wing. The downstream displacement of the trailing edge generated vortices, limits the downforce generated by those vortices also regularising the shedding cycle that appears to be much more organised when the flaplet is activated. A similar study has also been carried out for the dynamic case. We have analysed the effects produced by the presence of an elastically mounted flap on the transient behaviour of the flow fields. For a specific ramp-up manoeuvre characterised by a reduced frequency slower the shedding one, it is found that it is possible to design flaps that limit the severity of the dynamic stall breakdown. In particular, it is possible to increase the value of the lift overshoot and to smooth its abrupt decay in time. A detailed analysis on the modification of the unsteady vorticity field due to the flap-flow interaction during the ramp-up motion is also provided to explain the physical mechanism that lead to more benign aerodynamic response.

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20

Wilks, Brett Landon Burkhalter Johnny Evans. "Aerodynamics of wrap-around fins in supersonic flow." Auburn, Ala., 2005. http://repo.lib.auburn.edu/2005%20Fall/Thesis/WILKS_BRETT_54.pdf.

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21

Rabe, Angela C. "Effectiveness of a Serpentine Inlet Duct Flow Control Scheme at Design and Off-Design Simulated Flight Conditions." Diss., Virginia Tech, 2003. http://hdl.handle.net/10919/28653.

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An experimental investigation was conducted in a static ground test facility to determine the flow quality of a serpentine inlet duct incorporating active flow control for several simulated flight conditions. The total pressure distortion at the aerodynamic interface plane (AIP) was then used to predict the resulting stability for a compression system. This study was conducted using a model of a compact, low observable, engine inlet duct developed by Lockheed Martin. A flow control technique using air injection through microjets at 1% of the inlet mass flow rate was developed by Lockheed Martin to improve the quality of the flow exiting the inlet duct. Both the inlet duct and the flow control technique were examined at cruise condition and off-design simulated flight conditions (angle of attack and asymmetric distortion). All of the experimental tests were run at an inlet throat Mach number of 0.55 and a resulting Reynolds number of 1.76*105 based on the hydraulic diameter at the inlet throat. For each of the flight conditions tested, the flow control scheme was found to improve the flow uniformity and reduce the inlet distortion at the AIP. For simulated cruise condition, the total pressure recovery was improved by ~2% with the addition of flow control. For the off-design conditions of angle of attack and asymmetric distortion, the total pressure recovery was improved by 1.5% and 2% respectively. All flight conditions tested showed a reduction in circumferential distortion intensity with flow control. The cruise condition case showed reduced maximum circumferential distortion of 70% with the addition of flow control. A reduction in maximum circumferential distortion of 40% occurred for the angle of attack case with flow control, and 30% for the asymmetric distortion case with flow control. The inlet total pressure distortion was used to predict the changes in stability margin of a compression system due to design and off-design flight conditions and the improvement of the stability margin with the addition of flow control. A parallel compressor model (DYNTECC) was utilized to predict changes in the stability margin of a representative compression system (NASA Stage 35). Without flow control, all three cases show similar reduced stability margins on the order of 30% of the original stability margin for NASA Stage 35 at 70% corrected rotor speed. With the addition of flow control, the cruise condition tested improved the stability margin to 80% of the original value while the off-design conditions recover to 60% of the original margin. Overall, the flow control has been found to be extremely beneficial in improving the operating range of a compression system for the same inlet duct without flow control.
Ph. D.

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22

Coutley, Raymond L. "Numerical studies of compressible flow over a double-delta wing at high angle of attack." Thesis, Monterey, California. Naval Postgraduate School, 1990. http://hdl.handle.net/10945/30688.

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Approved for public release, distribution is unlimited
The objective of this work is the investigation of vortical flows at high angles of attack using numerical techniques. The first step for a successful application of a numerical technique, such as fimite difference or finite volume, is the generation of a computational mesh which can capture adequately and accurately the important physics of the flow. Therefore, the first part of this work deals with the grid generation over a double-delta wing and the second part deals with the visualization of the computed flow field over the double-delta wing at different angles of attack. The surface geometry of the double-delta wing is defined algebraically. The developed surface grid generator provides flexibility in distributing the surface points along the axial and circumferential directions. The hyperbolic grid generation method is chosen for the field grid generation and both cylindrical and spherical grids are constructed. The computed low speed (M = 0.2) flow results at different angles of attack over the double-delta wing are visualized. Important flow characteristics of the leeward side flow field are discussed while the development of vortex interaction, occurrence and progression of vortex breakdown as the angle of attack increases is demonstrated. The computed results at different fixed angles of attack are presented.

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23

Lego, Zachary Michael. "Analysis of High Angle of Attack Maneuvers to Enhance Understanding of the Aerodynamics of Perching." University of Dayton / OhioLINK, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1355101333.

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24

Nadeau, Yvan. "A comparative non-linear analytical and numerical irrotational analysis of aerofoils at high angle of attack /." Thesis, McGill University, 1989. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=55626.

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25

Lung, Ming-Hung. "Flowfield measurements in the vortex wake of a missile at high angle of attack in turbulence." Thesis, Monterey, California. Naval Postgraduate School, 1988. http://hdl.handle.net/10945/23235.

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Approved for public release; distribution is unlimited
The flowfield downstream of a vertically-launched surface-to-air missile model at an angle of attack of 50° and a Reynolds number of 1.1 x 10(5) was investigated in a wind tunnel of the Naval Postgraduate School. The goal of this thesis is to experimentally validate the pressure measurement system for flowfield variables with elevated levels of turbulence; to determine the location and intensity of the asymmetric vortices in the wake of the VLSAM model at a raised level of freestream turbulence; and to display the asymmetric vortices by velocity mapping and pressure contours. The purpose is to correlate the results with the force measurements of Rabang to provide a greater understanding of the vortex flowfield. The body-only configuration was tested. Two flowfield conditions were treated: the nominal ambient wind tunnel condition, and a condition with grid generated turbulence of 3.8% turbulence intensity and a dissipation length scale of 1.7 inches. The following conclusions were reached: 1) The relative strengths of the asymmetric vortices can be noted by the sharp spike shape in the ambient condition; this condition becomes diffused and becomes fatter in the turbulent condition; 2) The right side vortex has greater strength than the left side one as seen by the diffusion in the total pressure coefficient and static pressure coefficient contours with and without a turbulent condition; 3) an increase in turbulence intensity tends to reduce the strength of the asymmetric nose-generated vortices; also pushes the two asymmetric vortices closer together; 4) and crossflow velocities were examined and were found to indicate the behavior denoted by the pressure contours.
http://archive.org/details/flowfieldmeasure00lung
Lieutenant, Republic of China Navy

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26

Takahama, Morio, Noboru Sakamoto, and Yuhei Yamato. "Attitude Stabilization of an Aircraft via Nonlinear Optimal Control Based on Aerodynamic Data." Institute of Electrical and Electronics Engineers, 2009. http://hdl.handle.net/2237/14420.

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27

Contreras, Daimer Mauthsud Leovan Ospina. "Angle of attack impact in the aerothermodynamics of a hypersonic vehicle with surface discontinuity-like a cavity." Instituto Nacional de Pesquisas Espaciais (INPE), 2017. http://urlib.net/sid.inpe.br/mtc-m21b/2017/04.12.00.58.

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The study described in this dissertation was undertaken with the purpose to investigate the impact of discontinuities present on the surface of hypersonic space vehicles. With this perspective in mind, computational simulations of a non-reacting rarefied hypersonic flow over a flat plate with a cavity have been performed by using the Direct Simulation Monte Carlo method. Simulations provided a comprehensive description about the nature of the flowfield structure and the aerodynamic surface properties on the cavity resulting from changes in the length-to-depth (L/H) ratio and changes in the angle of attack of the oncoming flow. A detailed description of the flowfield properties (velocity, density, pressure and temperature) and aerodynamics surface properties (number flux, heat transfer, pressure and skin friction) were obtained by a numerical method that properly account for non-equilibrium effects in the transition flow regime. Results for a cavity defined by L/H ratio of 1, 2, 3 and 4, and flow with angle of attack of 10, 15 and 20 degrees, were compared to those of a flat plate without a cavity with zero-degree angle of incidence and with a flat plate at incidence. The analysis showed that the flow topology inside the cavity, composed by recirculation regions, depended on the L/H ratio as well as on the angle of attack, for the conditions investigated. For L/H < 3 a single vortex core was formed, and filled entirely the cavity. In contrast, for L/H of 3 and 4, two vortices were formed inside the cavity, at the vicinity of the backward and forward faces. The analysis also showed that, for the L/H = 4 case, the flow topology inside the cavity corresponds to that of a ${''}$closed cavity${''}$ in the continuum flow regime for 10-degree angle of incidence, and similar to an open cavity for the others angles of attack investigated. In addition, it was found that the maximum values for the heat transfer, pressure and skin friction coefficients inside the cavity took place on the cavity forward face. It was also found that, maximum values for heat transfer coefficient inside the cavities increased with increasing the angle of attack $\alpha$. However, it was observed that these maximum values are smaller than those observed in a flat-plate without a cavity for the corresponding angle of attack. Consequently, in terms of pressure, the presence of the cavity on the vehicle surface can not be ignored in the vehicle design.
O estudo descrito nesta dissertação foi realizado com o propósito de investigar o impacto de descontinuidades presentes na superfície de veículos espaciais hipersônicos. Em busca deste propósito, simulações computacionais de um escoamento hipersônico rarefeito não-reativo sobre uma cavidade foram realizadas usando-se o método Direct Simulation Monte Carlo. As simulações forneceram informações detalhadas sobre a natureza da estrutura do escoamento, propriedades primárias e propriedades aerodinâmicas, em função de mudanças na razão comprimento-profundidade (L/H) da cavidade, e mudanças no ângulo de ataque do escoamento incidindo sobre a cavidade. Uma descrição detalhada, das propriedades primárias (velocidade, massa específica, pressão e temperatura) e das quantidades aerodinâmica na superfície (transferência de calor, pressão e atrito), foi obtida por um método numérico que leva em conta adequadamente os efeitos de não-equilíbrio no regime de transição. Os resultados, para cavidades definidas por L/H de 1, 2, 3 e 4, com ângulos de ataque do escoamento de 10, 15 e 20 graus, foram comparados com os de uma placa plana sem/com a presença de cavidade sem/com incidência. A análise mostrou que a topologia do escoamento dentro da cavidade, composta por regiões de recirculação,dependeu da razão L/H bem como do ângulo de ataque do escoamento, para as condições investigadas. Para L/H < 3, observou-se a formação de um único vórtice ocupando inteiramente a cavidade. Para cavidade com L/H =3 e 4, dois vórtices foram formados dentro da cavidade, nas vizinhanças das faces a montante e a jusante da cavidade. A análise também mostrou que, para uma cavidade com L/H = 4 e 10 graus de incidência, a estrutura do escoamento dentro da cavidade correspondeu aquela de uma cavidade fechada , conforme definido para um escoamento no regime do contínuo. Por outro lado, para L/H = 4 e maiores ângulos de incidência, a estrutura do escoamento correspondeu aquela de uma cavidade aberta , para os ângulos de ataque investigados. Outrossim, verificou-se que os valores máximos para os coeficientes de transferência de calor, pressão e coeficiente de atrito ocorreram na superfície a montante do escoamento dentro da cavidade. Verificou-se também que, os valores máximos para o coeficiente de transferência de calor dentro da cavidade aumentaram com o aumento do ângulo de ataque $\alpha$. Todavia, esses valores máximos foram menores do que aqueles observados sobre uma placa plana sem cavidade com incidência. Como resultado, em termos de pressão, a presença da cavidade sobre a superfície do veículo não pode ser ignorada no projeto do veículo.

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Atkinson, Michael D. "Control of Hypersonic High Angle-Of-Attack Re-Entry Flow Using a Semi-Empirical Plasma Actuator Model." University of Dayton / OhioLINK, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1335283726.

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29

Whale, James Callum Andrew. "How to make a tuna burst : the role of angle of attack in the production of thrust." Thesis, University of British Columbia, 2016. http://hdl.handle.net/2429/58577.

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Tuna—along with whales and lamnid sharks—utilise thunniform locomotion, a mode of swimming that optimises efficiency at high speed and isolates thrust production to the caudal fin. Thunniform performance is controlled by adjustments in the way the caudal fin interacts with fluid flow which, in turn, determines thrust and efficiency. The effect of tail motion on performance provides insight into the link between locomotor muscle biomechanics and hydrodynamics; these insights can be used to mimic and optimise animal motion in a robotic context. This study focuses on how the maximum angle of attack (α_max), contributes to tuna cruising and bursting, and the corresponding effects on fluid flow. I hypothesise that cruising tuna do not adjust α_max to modulate thrust but instead vary amplitude via Strouhal number. I also hypothesise that α_max affects thrust by changes in vorticity shed by the tail. To study these phenomena, I constructed a tuna tail model 3-D printed from CT scan data of a tuna tail. I then oscillated this model in a water tunnel across a range of biologically relevant motions. I calculated thrust and efficiency from direct measurements of force and torque and then used ink-flow visualisation and particle image velocimetry to reveal the resulting flow structures. The results indicate that the efficiency optimum of α_max peaks around 15° with the thrust optimum beyond 30°. Mechanistically, an increase of α_max increases the magnitude of the resultant force but angles it to the side, increasing the amount of wasted lateral energy. Increasing α_max increases the size and strength of shed vortices eventually causing shedding of an additional leading edge vortex at midstroke. These results, paired with red muscle work loop data, suggest that during cruise the α_max undergoes minimal variation, and suggest that in order to take advantage of the additional thrust that high values of α_max provide, white burst muscles need to advance peak force timing. In addition to contributing to a better understanding of the hydrodynamics of swimming and the associated musculature, these results also offer insight into the field of biomimetics and the construction of fish-mimicking robots such as AUVs.
Science, Faculty of
Zoology, Department of
Graduate

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30

Lopera, Javier. "Aerodynamic Control of Slender Bodies from Low to High Angles of Attack through Flow Manipulation." Connect to Online Resource-OhioLINK, 2007. http://www.ohiolink.edu/etd/view.cgi?acc_num=toledo1177504352.

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31

Alkhozam, Abdullah M. "Interaction, bursting and dynamic control of vortices of a cropped double-delta wing at high angle of attack." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1994. http://www.dtic.mil/cgi-bin/GetTRDoc?AD=ADA283656&Location=U2&doc=GetTRDoc.pdf.

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Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, March 1994.
Thesis advisor(s): S. K. Hebbar, M. F. Platzer. "March 1994." Cover title: Interaction, ... and control of vortices of a cropped ... Includes bibliographical references. Also available online.

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32

Lewis, Daniel Joseph. "Tip clearance and angle of attack effects upon the unsteady response of a vibrating flat plate in crossflow /." This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-06112009-063924/.

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33

Torres, Luis Carlos Roldan. "Angle of attack effect in the aerothermodynamics of a hypersonic vehicle with a surface discontinuity of gap type." Instituto Nacional de Pesquisas Espaciais (INPE), 2017. http://urlib.net/sid.inpe.br/mtc-m21b/2017/05.23.23.55.

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The development of hypersonic vehicles has become a topic of interest in recent years, since has made it possible to reach inaccessible places such as orbital flights. The construction of these vehicles must be made with specials materials, and must have an efficient aerodynamic shape to withstand high speeds, high temperatures and significant pressure changes. The study described in this dissertation was undertaken with the objective to investigate the impact of discontinuities present on the surface of hypersonic space vehicles. In pursuit of this goal, computational simulations of a low-density hypersonic flow over a flat plate with a gap has been performed by using the Direct Simulation Monte Carlo method. The simulations provided information about the nature of the flowfield structure and the aerodynamic surface properties on the gap resulting from variations in the length-to-depth (L/H) ratio and variations in the angle of attack. A description of the flowfield properties, such as velocity, density, pressure and temperature, and aerodynamics surface quantities, such as, number flux, heat transfer, pressure and skin friction, were obtained by a numerical method that properly account for non-equilibrium effects in the transition flow regime. Results for a gap defined by L/H ratio of 1, 1/2, 1/3 and 1/4, and flow with angle of attack of 10, 15 and 20 degrees, were compared to those of a flat plate without a gap with zero-degree angle of incidence. The analysis showed that the flow topology inside the gap with incidence is slightly different from that for zero-degree angle of incidence for the L/H ratio investigated. It was found that the maximum values for the heat transfer, pressure and skin friction coefficients inside the gap took place on the gap forward face. It was also found that, maximum values for heat transfer coefficient inside the gaps increased with increasing the angle of attack $\alpha$. Nevertheless, it was observed that these maximum values are smaller than those observed in a flat-plate without a gap for the corresponding angle of attack. As a result, in terms of pressure, the presence of the gap on the vehicle surface can not be ignored in the vehicle design.
O desenvolvimento de veículos hipersônicos tem se tornado um tema de interesse nos últimos anos, considerando-se a possibilidade de se chegar com tais veículos a locais até então inacessíveis como os voos orbitais. A construção desses veículos exige materiais especiais e deve apresentar uma forma aerodinâmica eficiente para resistir altas velocidades além de temperaturas elevadas e mudanças de pressão significativas. O estudo descrito nesta dissertação foi realizado com o objetivo de investigar o impacto de descontinuidades presentes na superfície de veículos espaciais hipersônicos. Em busca deste objetivo, simulações computacionais de um escoamento hipersônico rarefeito sobre uma placa plana, foi realizada usando-se o método Direct Simulation Monte Carlo. As simulações forneceram informações sobre a natureza da estrutura do escoamento, propriedades primarias e propriedades aerodinâmicas, devido a variações na razão comprimento-profundidade (L/H), e variações no ângulo de ataque. Uma descrição das propriedades primarias, tais como velocidade, massa específica, pressão e temperatura, e das quantidades aerodinâmica, tais como transferência de calor, pressão e atrito na superfície, foi obtida por um método numérico que leva em conta os efeitos de não-equilíbrio no regime de transição. Os resultados para um filete definido por uma razão L/H de 1, 1/2, 1/3 e 1/4, e com ângulo de ataque do escoamento de 10, 15 e 20 graus, foram comparados com os de uma placa plana sem a presença de um filete. A análise mostrou que a estrutura do escoamento dentro do filete com ângulo de ataque é ligeiramente diferente daquela com zero grau de incidência para cada razão L/H investigada. Verificou-se que os valores máximos para os coeficientes de transferência de calor, pressão e coeficiente de atrito ocorreram na superfície a montante do escoamento dentro do filete. Verificou-se também que, os valores máximos para o coeficiente de transferência de calor dentro do filete aumentaram com o aumento do ângulo de ataque $\alpha$. Como resultado, em termos de pressão, a presença do filete sobre a superfície do veículo não pode ser ignorada no projeto do veículo.

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34

Lewis, Daniel Russell. "Tip clearance and angle of attack effects upon the unsteady response of a vibrating flat plate in crossflow." Thesis, Virginia Tech, 1993. http://hdl.handle.net/10919/43198.

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The influence of tip clearance and angle of attack upon the mid-span unsteady pressure response of a vibrating flat plate was investigated experimentally. Unsteady pressure measurements were taken for a variety of incidence angles, vibration frequencies and tip clearances over a Mach number range of 0.2 to 0.6.

It was found that changes in tip clearance had an effect on measured pressure fluctuations at higher angles of attack and larger Mach numbers. It was also observed that the amplitude of the unsteady pressure increased as the incidence angle was increased.

The plate was mechanically induced to oscillate in translation, simulating the flISt bending mode. Averaged Fast Fourier Transforms were used to determine pressure oscillation amplitudes and phase lags with respect to the plate motion.

Master of Science

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35

東, 大輔, Daisuke AZUMA, 佳朗中村, and Yoshiaki NAKAMURA. "前縁回転/後縁ジェットハイブリッド法によるデルタ翼揚力増加." 日本航空宇宙学会, 2006. http://hdl.handle.net/2237/13878.

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36

Najarzadegan, Farshid. "Circulation Dependence of the Interaction Between a Wing-Tip Vortex and Turbulence." UKnowledge, 2019. https://uknowledge.uky.edu/me_etds/145.

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Vortices are present in many fluid flows and depending on the context they may be either beneficial or harmful for different systems or processes. Planar particle image velocimetry was used to examine the vortex evolution and its decay under different turbulence intensities and vortex circulation. The vortex decayed faster in the presence of high turbulence intensity. Vortex trajectories were impacted by turbulence intensity and vortex strength. Trajectories with no turbulence intensity had less variation. The vortex wandering amplitude decreased with growth of vortex strength. The vortex decay was confined to the core of the vortex, with the tangential velocity at large radial distances from the vortex center being relatively constant in time. The vortex core radius had a greater rate of growth with the low turbulence intensity and lower angle of attack. The amplitude of fluctuation of the core circulation increased for the higher turbulence intensity and weaker vortex.

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37

Schaeffler, Norman Walter. "All The King's Horses: The Delta Wing Leading-Edge Vortex System Undergoing Vortex Breakdown: A Contribution to its characterization and Control under Dynamic Conditions." Diss., Virginia Tech, 1998. http://hdl.handle.net/10919/30454.

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The quality of the flow over a 75 degree-sweep delta wing was documented for steady angles of attack and during dynamic maneuvers with and without the use of two control surfaces. The three-dimensional velocity field over a delta wing at a steady angle of attack of 38 degrees and Reynolds number of 72,000 was mapped out using laser-Doppler velocimetry over one side of the wing. The three-dimensional streamline and vortex line distributions were visualized. Isosurfaces of vorticity, planar distributions of helicity and all three vorticity components, and the indicator of the stability of the core were studied and compared to see which indicated breakdown first. Visualization of the streamlines and vortex lines near the core of the vortex indicate that the core has a strong inviscid character, and hence Reynolds number independence, upstream of breakdown, with viscous effects becoming more important downstream of the breakdown location. The effect of cavity flaps on the flow over a delta wing was documented for steady angles of attack in the range 28 degrees to 42 degrees by flow visualization and surface pressure measurements at a Reynolds number of 470,000 and 1,000,000, respectfully. It was found that the cavity flaps postpone the occurrence of vortex breakdown to higher angles of attack than can be realized by the basic delta wing. The effect of continuously deployed cavity flaps during a dynamic pitch-up maneuver of a delta wing on the surface pressure distribution were recorded for a reduced frequency of 0.0089 and a Reynolds number of 1,300,000. The effect of deploying a set of cavity flaps during a dynamic pitch-up maneuver on the surface pressure distribution was recorded for a reduced frequency of 0.0089 and a Reynolds number of 1,300,000 and 187,000. The active deployment of the cavity flaps was shown to have a short-lived beneficial effect on the surface pressure distribution. The effect on the surface pressure distribution of the varying the reduced frequency at constant Reynolds number for a plain delta wing was documented in the reduced frequency range of 0.0089 to 0.0267. The effect of the active deployment of an apex flap during a pitch-up maneuver on the surface pressure distribution at Reynolds numbers of 532,000, 1,000,000, and 1,390,000 were documented with reduced frequencies of 0.0053 to 0.0114 with flap deployment locations in the range of 21Â° to 36Â° . The apex flap deployment was found to have a beneficial effect on the surface pressure distribution during the maneuver and in the post-stall regime after the maneuver is completed.
Ph. D.

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38

Webb, Charles. "Separation and Vorticity Transport in Massively-Unsteady Low Reynolds Number Flows." Wright State University / OhioLINK, 2009. http://rave.ohiolink.edu/etdc/view?acc_num=wright1244864717.

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39

Findlay, David Bruce. "A numerical study of aircraft empennage buffet." Diss., Georgia Institute of Technology, 1999. http://hdl.handle.net/1853/10926.

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40

De, Oliveira Neto Pedro Jose. "An Investigation of Unsteady Aerodynamic Multi-axis State-Space Formulations as a Tool for Wing Rock Representation." Diss., Virginia Tech, 2007. http://hdl.handle.net/10919/29600.

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The objective of the present research is to investigate unsteady aerodynamic models with state equation representations that are valid up to the high angle of attack regime with the purpose of evaluating them as computationally affordable models that can be used in conjunction with the equations of motion to simulate wing rock. The unsteady aerodynamic models with state equation representations investigated are functional approaches to modeling aerodynamic phenomena, not directly derived from the physical principles of the problem. They are thought to have advantages with respect to the physical modeling methods mainly because of the lower computational cost involved in the calculations. The unsteady aerodynamic multi-axis models with state equation representations investigated in this report assume the decomposition of the airplane into lifting surfaces or panels that have their particular aerodynamic force coefficients modeled as dynamic state-space models. These coefficients are summed up to find the total aircraft force coefficients. The products of the panel force coefficients and their moment arms with reference to a given axis are summed up to find the global aircraft moment coefficients. Two proposed variations of the state space representation of the basic unsteady aerodynamic model are identified using experimental aerodynamic data available in the open literature for slender delta wings, and tested in order to investigate their ability to represent the wing rock phenomenon. The identifications for the second proposed formulation are found to match the experimental data well. The simulations revealed that even though it was constructed with scarce data, the model presented the expected qualitative behavior and that the concept is able to simulate wing rock.
Ph. D.

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41

Reuter, William H. Howard Richard M. Hobson Garth V. Buning Pieter G. "Flowfield computations over the space shuttle Orbiter with a proposed canard at a Mach number of 5.8 and 50 degrees angle of attack /." Monterey, Calif. : Springfield, Va. : Naval Postgraduate School; Available from the National Technical Information Service, 1993. http://handle.dtic.mil/100.2/ADA258058.

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Thesis (Degree of Aeronautical and Astronautical Engineer) Naval Postgraduate School, June 1993.
Thesis advisors, Richard M. Howard, Garth V. Hobson and Pieter G. Buning. AD-A258 058. Includes bibliographical references. Also available online.

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42

Reuter, William H. IV. "Flowfield computations over the space shuttle Orbiter with a proposed canard at a Mach number of 5.8 and 50 degrees angle of attack." Thesis, Monterey, California. Naval Postgraduate School, 1993. http://hdl.handle.net/10945/39837.

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43

Mogili, Prasad. "RANS and DES computations for a three-dimensional wing with ice accretion." Master's thesis, Mississippi State : Mississippi State University, 2004. http://library.msstate.edu/etd/show.asp?etd=etd-07102004-145100.

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44

Lapinskas, Vytautas. "Sklandytuvo atakos ir slydimo kampų matavimo metodų tyrimas." Master's thesis, Lithuanian Academic Libraries Network (LABT), 2011. http://vddb.laba.lt/obj/LT-eLABa-0001:E.02~2011~D_20110615_175706-86724.

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Baigiamajame magistro darbe atliekamas sklandytuvo atakos ir slydimo kampo matavimo metodų tyrimas. Pirmoje darbo dalyje apžvelgiami atakos kampo matuokliai: virvutė, pritvirtinta ant stiklinio gaubto, atakos – slydimo kampo matuoklis su vėjarodžiu ir Pitoto vamzdelio tipo daviklis. Davikliai palyginami, aprašomi jų privalumai ir trūkumai lyginant su kitais davikliais. Antroje dalyje aprašomi alfa ir beta kampų matavimo metodai: matavimas vamzdelio tipo davikliu ir metodas, kai nenaudojami specialūs atakos, slydimo kampo davikliai. Toliau apžvelgiami veiksniai, turintys įtakos matavimo tikslumui. Pateikiamos kelių vamzdelio tipo daviklių kalibravimo kreivės. Paskutinėje dalyje programa Matlab kuriamas matematinis-dinaminis sklandytuvo modelis. Modeliu, pagal nustatytas sąlygas, skaičiuojami atakos ir slydimo kampai, analizuojami grafikai.
The thesis made the glider’s attack and slip angle measurement methods for the investigation. The first part gives an overview of measuring devices of angle of attack and slip angle: The side string, attached to the side of the canopy, vane mounted AOA sensor, Pitot-tube type sensor. The sensors are compared, describes their advantages and disadvantages compared with other sensors. The second part describes the alpha and beta angle measurement methods: measurement with the tube-type sensor, and the method without using the specific attack, slip angle sensors. The following gives an overview of factors affecting the measurement accuracy. Several tube-type sensor calibration curves are presented. The last part of thesis presents development of mathematical – dynamic model of the glider using Matlab software. The model calculates the angle of attack and slip using set conditions of flight.

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Gomes, Lara Elena. "Comparação entre forças propulsivas efetivas calculadas e medida durante um palmateio de sustentação." reponame:Biblioteca Digital de Teses e Dissertações da UFRGS, 2010. http://hdl.handle.net/10183/27677.

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A força propulsiva gerada durante o palmateio é resultado do somatório das forças de arrasto e de sustentação, sendo que a componente que atua na direção do movimento desejado é igual à força propulsiva efetiva. Essas forças podem ser estimadas a partir de equações hidrodinâmicas, porém essas equações não consideram todos os mecanismos que contribuem para a propulsão. Dessa forma, o objetivo geral do presente estudo foi comparar a força propulsiva efetiva calculada a partir das equações hidrodinâmicas e a força propulsiva efetiva medida durante o palmateio de sustentação (na posição vertical, de cabeça para cima) em cada fase do palmateio. Para isso, uma praticante de nado sincronizado realizou palmateio na posição vertical de cabeça para cima durante 15 segundos, enquanto que dados cinemáticos e cinéticos foram obtidos por viodeogrametria 3D e dinamometria respectivamente. A análise gráfica de Bland e Altman foi usada para comparar as forças propulsivas efetivas medida e calculada durante o palmateio. As forças propulsivas efetivas calculada e medida foram diferentes, sendo a medida maior que a calculada. Ainda, os resultados indicaram que o palmateio executado não foi simétrico, isto é, a orientação e a força propulsiva entre a mão direita e a esquerda foram diferentes. Portanto, o achado do presente trabalho destaca a importância de mecanismos instáveis para a propulsão durante o palmateio, já que as forças estimadas por meio das equações hidrodinâmicas apresentaram resultados inferiores, sendo isso observado ao longo de todo o palmateio.
Propulsive force generated during sculling motion results from drag and lift propulsive forces, and the component acting in the direction of motion is the effective propulsive force. These forces may be calculated using hydrodynamic equations, but these equations do not consider all mechanisms that contribute to the propulsion. Thus, the main purpose of this study was to compare the calculated effective propulsive force using the hydrodynamic equations and the measured effective propulsive force during a support sculling motion (vertical position with the head above the water‟s surface) in each phase of sculling. For this, a practitioner of synchronized swimming performed sculling motion in a vertical position with the head above the water‟s surface during 15 seconds, while kinematic and kinetic data were obtained by 3D videogrammetry and dynamometry respectively. Graphical techniques from Bland and Altman were used to compare the measured effective propulsive force and calculated effective propulsive force during sculling motion. The calculated effective propulsive force and the measured effective propulsive force were different, the measured being greater than the calculated. Moreover, the results indicated sculling motion performed was not symmetric, that is, the orientation and propulsive forces between the right and left hands were different. Therefore, the result of this study highlights the importance of the unsteady mechanisms for the propulsion during sculling motion, because the calculated forces using the hydrodynamic equations presented low values throughout the sculling motion.

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46

Salemi, Leonardo da Costa, and Leonardo da Costa Salemi. "Numerical Investigation of Hypersonic Conical Boundary-Layer Stability Including High-Enthalpy and Three-Dimensional Effects." Diss., The University of Arizona, 2016. http://hdl.handle.net/10150/621854.

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The spatial stability of hypersonic conical boundary layers is investigated utilizing different numerical techniques. First, the development and verification of a Linearized Compressible Navier-Stokes solver (LinCS) is presented, followed by an investigation of different effects that affect the stability of the flow in free-flight/ground tests, such as: high-enthalpy effects, wall-temperature ratio, and three-dimensionality (i.e. angle-of-attack). A temporally/spatially high-order of accuracy parallelized Linearized Compressible Navier-Stokes solver in disturbance formulation was developed, verified and employed in stability investigations. Herein, the solver was applied and verified against LST, PSE and DNS, for different hypersonic boundary-layer flows over several geometries (e.g. flat plate - M=5.35 & 10; straight cone - M=5.32, 6 & 7.95; flared cone - M=6; straight cone at AoA = 6 deg - M=6). The stability of a high-enthalpy flow was investigated utilizing LST, LinCS and DNS of the experiments performed for a 5 deg sharp cone in the T5 tunnel at Caltech. The results from axisymmetric and 3D wave-packet investigations in the linear, weakly, and strongly nonlinear regimes using DNS are presented. High-order spectral analysis was employed in order to elucidate the presence of nonlinear couplings, and the fundamental breakdown of second mode waves was investigated using parametric studies. The three-dimensionality of the flow over the Purdue 7 deg sharp cone at M=6 and AoA =6 deg was also investigated. The development of the crossflow instability was investigated utilizing suction/blowing at the wall in the LinCS/DNS framework. Results show good agreement with previous computational investigations, and that the proper basic flow computation/formation of the vortices is very sensitive to grid resolution.

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47

Morrison, Thomas M. "THE USE OF TELEMETRY DATA IN AN AIR DATA SYSTEM." International Foundation for Telemetering, 2006. http://hdl.handle.net/10150/604135.

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ITC/USA 2006 Conference Proceedings / The Forty-Second Annual International Telemetering Conference and Technical Exhibition / October 23-26, 2006 / Town and Country Resort & Convention Center, San Diego, California
Telemetry data are usually collected for analysis at some later time and can be monitored to follow the progress of a test. In the case of an Air Data System the signals from the sensors are sent to a computer that calculates the air data parameters for use on multiple LabView-generated displays, as well as to the Data Acquisition System. The readouts on the multiple displays need to be real-time so they are useful to the flight crew. Equations that control the different air data values are determined by what telemetry data are available and the preference of those doing the test planning. These systems need to display the information in a format useful to the flight crew and be reliable.

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48

Le, Moigne Yann. "Adaptive Mesh Refinement and Simulations of Unsteady Delta-Wing Aerodynamics." Doctoral thesis, KTH, Aeronautical and Vehicle Engineering, 2004. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3786.

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This thesis deals with Computational Fluid Dynamics (CFD)simulations of the flow around delta wings at high angles ofattack. These triangular wings, mainly used in militaryaircraft designs, experience the formation of two vortices ontheir lee-side at large angles of attack. The simulation ofthis vortical flow by solving the Navier-Stokes equations isthe subject of this thesis. The purpose of the work is toimprove the understanding of this flow and contribute to thedesign of such a wing by developing methods that enable moreaccurate and efficient CFD simulations.

Simulations of the formation, burst and disappearance of thevortices while the angle of attack is changing are presented.The structured flow solver NSMB has been used to get thetime-dependent solutions of the flow. Both viscous and inviscidresults of a 70°-swept delta wing pitching in anoscillatory motion are reported. The creation of the dynamiclift and the hysteresis observed in the history of theaerodynamic forces are well reproduced.

The second part of the thesis is focusing on automatic meshrefinement and its influence on simulations of the delta wingleading-edge vortices. All the simulations to assess the gridquality are inviscid computations performed with theunstructured flow solver EDGE. A first study reports on theeffects of refining thewake of the delta wing. A70°-swept delta wing at a Mach number of 0.2 and an angleof attack of 27° where vortex breakdown is present abovethe wing, is used as testcase. The results show a strongdependence on the refinement, particularly the vortex breakdownposition, which leads to the conclusion that the wake should berefined at least partly. Using this information, a grid for thewing in the wind tunnel is created in order to assess theinfluence of the tunnel walls. Three sensors for automatic meshrefinement of vortical flows are presented. Two are based onflow variables (production of entropy and ratio of totalpressures) while the third one requires an eigenvalue analysisof the tensor of the velocity gradients in order to capture theposition of the vortices in the flow. These three vortexsensors are successfully used for the simulation of the same70° delta wing at an angle of attack of 20°. Acomparison of the sensors reveals the more local property ofthe third one based on the eigenvalue analysis. This lattertechnique is applied to the simulation of the wake of a deltawing at an angle of attack of 20°. The simulations on ahighly refined mesh show that the vortex sheet shed from thetrailing-edge rolls up into a vortex that interacts with theleading-edge vortex. Finally the vortex-detection technique isused to refine the grid around a Saab Aerosystems UnmannedCombat Air Vehicle (UCAV) configuration and its flight dynamicscharacteristics are investigated.

Key words:delta wing, high angle of attack, vortex,pitching, mesh refinement, UCAV, vortex sensor, tensor ofvelocity gradients.

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49

Mitori, Tiffany Leilani. "Flight and Stability of a Laser Inertial Fusion Energy Target in the Drift Region between Injection and the Reaction Chamber with Computational Fluid Dynamics." DigitalCommons@CalPoly, 2014. https://digitalcommons.calpoly.edu/theses/1154.

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Abstract:

A Laser Inertial Fusion Energy (LIFE) target’s flight through a low Reynolds number and high Mach number regime was analyzed with computational fluid dynamics software. This regime consisted of xenon gas at 1,050 K and approximately 6,670 Pa. Simulations with similar flow conditions were performed over a sphere and compared with experimental data and published correlations for validation purposes. Transient considerations of the developing flow around the target were explored. Simulations of the target at different velocities were used to determine correlations for the drag coefficient and Nusselt number as functions of the Reynolds number. Simulations with different target angles of attack were used to determine the aerodynamic coefficients of drag, lift, Magnus moment, and overturning moment as well as target stability. The drag force, lift force, and overturning moment changed minimally with spin. Above an angle of attack of 15°, the overturning moment would be destabilizing. At angles of attack less than 15°, the overturning moment would tend to decrease the target’s angle of attack, indicating the lack of a need for spin for stability at these small angles. This stabilizing moment would cause the target to move in a mildly damped oscillation about the axis parallel to the free-stream velocity vector through the target’s center of gravity.

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50

Horpatzká, Michaela. "Systém pro haptickou odezvu a jeho spolehlivost." Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2019. http://www.nusl.cz/ntk/nusl-401539.

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Abstract:

This master thesis covers conceptual design of haptic feedback system and safety and reliability assessment. System is developed for aircraft categories UL2, EASA CS-LSA, EASA CS-VLA and EASA-CS-23. Conceptual design is divided into three parts. First is aimed on haptic system test with simulator. Second is conceptual design of haptic system for in-flight test. Third part contains safety and reliability assessment and summarizes problems of haptic system installation in real aircraft.

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Dissertations / Theses on the topic 'Attack angle'

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