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1
Oh, Gyeongtaek, Jongho Park, Jeongha Park, Hongju Lee, Youdan Kim, Sang-Joon Shin, Jaemyung Ahn, and Sangbum Cho. "Load relief control of launch vehicle using aerodynamic angle estimation." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, no. 8 (March 23, 2017): 1598–605. http://dx.doi.org/10.1177/0954410017699435.
Повний текст джерелаАнотація:
A nonlinear closed-loop load relief scheme is proposed to reduce the aerodynamic load during the ascent phase of a launch vehicle. The proposed controller is designed based on a back-stepping and sliding-mode control scheme with aerodynamic angle feedback. A hybrid load-relief strategy using the load relief scheme around the period of the maximum dynamic pressure and the traditional minimum-drift scheme during the other period is proposed. An aerodynamic angle estimator is also developed using a Kalman filter for the feedback of the load relief control. Numerical simulation is conducted to demonstrate the performance of the proposed strategy as well as the potential benefits.
2
Tang, Wei, and Bi-Feng Song. "Transitional flight equilibrium and performance study for the X-NMRL tail-sitter VTOL MAV." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 8 (August 19, 2018): 3056–77. http://dx.doi.org/10.1177/0954410018794731.
Повний текст джерелаАнотація:
An investigation on transitional flight equilibrium, performance analysis and parameter impacts is conducted in a conversion corridor, based on the proposed X-NMRL tail-sitter Vertical Takeoff and Landing Micro Air Vehicles (VTOL MAVs). Dependent on a propulsion model, aerodynamic model and physical control model, a nonlinear mathematical transitional model of the vehicle dynamics was constructed with consideration of the velocity, angle of attack, thrust, control surface deflection and pitching angle. The momentum theory and estimation method are applied to simulated propeller slipstream effects on aerodynamics, and an aerodynamic model for all regions of angles of attack and velocities is built. The nonlinear indefinite high-order dynamic model is solved by the improved Newton iteration algorithm. The corridor of the pitching angle or flight-path angle to the velocity reveals that the boundaries are mainly governed by the stalling performance, full throttle thrust and zero thrust, respectively. The performance corridor indicates different performance parameter variations under different conditions of steady climbing, cruising and descending states. Additionally, the performance for a steady transitional strategy can be illustrated to some extent. In terms of the parameter impacts, the increasing max propulsive power, supplied voltage, and decreasing total weight can widen the transitional corridor effectively, and the changes in the aerodynamics will only move the boundaries toward the same direction. These results will benefit transitional vehicle designs and control designs.
3
Escobar-Ruiz, Alan G., Omar Lopez-Botello, Luis Reyes-Osorio, Patricia Zambrano-Robledo, Luis Amezquita-Brooks, and Octavio Garcia-Salazar. "Conceptual Design of an Unmanned Fixed-Wing Aerial Vehicle Based on Alternative Energy." International Journal of Aerospace Engineering 2019 (November 14, 2019): 1–13. http://dx.doi.org/10.1155/2019/8104927.
Повний текст джерелаАнотація:
This paper focuses on the aerodynamics and design of an unmanned aerial vehicle (UAV) based on solar cells as a main power source. The procedure includes three phases: the conceptual design, preliminary design, and a computational fluid dynamics analysis of the vehicle. One of the main disadvantages of an electric UAV is the flight time; in this sense, the challenge is to create an aerodynamic design that can increase the endurance of the UAV. In this research, the flight mission starts with the attempt of the vehicle design to get at the maximum altitude; then, the UAV starts to glide and battery charge recovery is achieved due to the solar cells. A conceptual design is used, and the aerodynamic analysis is focused on a UAV as a gliding vehicle, with the calculations starting with the estimation of weight and aerodynamics and finishing this stage with the best glide angle. In fact, the aerodynamic analysis is obtained for a preliminary design; this step involves the wing, fuselage, and empennage of the UAV. In order to achieve the preliminary design, an estimation of aerodynamic coefficients, along with computational fluid dynamics analysis, is performed.
4
Zhang, Jiaming, Qing Li, Nong Cheng, and Bin Liang. "Non-linear flight control for unmanned aerial vehicles using adaptive backstepping based on invariant manifolds." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 227, no. 1 (January 6, 2012): 33–44. http://dx.doi.org/10.1177/0954410011430027.
Повний текст джерелаАнотація:
A novel adaptive backstepping control scheme based on invariant manifolds for unmanned aerial vehicles in the presence of some uncertainties in the aerodynamic coefficients is presented in this article. This scheme is used for command tracking of the angle of attack, the sideslip angle, and the bank angle of the aircraft. The control law has a modular structure, which consists of a control module and a recently developed non-linear estimator. The estimator is based on invariant manifolds, which allows for prescribed dynamics to be assigned to the estimation error. The adaptive backstepping control law combined with the estimator covers the entire flight envelope and does not require accurate aerodynamic parameters. The stability of the whole closed-loop system is analyzed using the Lyapunov stability theory. The full six-degree-of-freedom non-linear model of a small unmanned aerial vehicle is used to demonstrate the effectiveness of the proposed control law. The numerical simulation result shows that this method can yield satisfying command tracking despite some unknown aerodynamic parameters.
5
Korsun, O. N., A. I. Daneko, P. A. Motlich, and M. H. Om. "Estimation of Angles of Attack and Sideslip of Unmanned Aerial Vehicle in the Absence of Aerodynamic Angle Sensors." Mekhatronika, Avtomatizatsiya, Upravlenie 23, no. 5 (May 6, 2022): 274–80. http://dx.doi.org/10.17587/mau.22.274-280.
Повний текст джерелаАнотація:
A method for estimating aerodynamic angles in the absence of appropriate sensors is proposed, using measurements of three projections of flight speed carried out by the navigation system and the values of the orientation angles. The relevance of the problem being solved is determined by the fact that on unmanned aerial vehicles (UAVs) sensors of aerodynamic angles, that is, angles of attack and slip, are often not installed due to restrictions on dimensions and mass. The proposed method is based on the joint use of mathematical models of aircraft motion, known from flight dynamics, and the theory of parametric identification of dynamic systems. The key factor ensuring the accuracy of the proposed method is the use of very accurate measurements of three UAV velocity projections performed by a satellite navigation system or an inertial navigation system with satellite correction. To account for the influence of wind, parametric identification of three projections of wind speed is provided. Another feature of the method is that instead of the missing aerodynamic angle sensors, it is proposed to use information about the aerodynamic coefficients of the lifting and lateral forces of the UAV. If these coefficients are known with errors, their values are also specified by identification methods. The dimension of the identification problem turns out to be low in the range of small and medium angles of attack when the aerodynamic dependencies are linear. The results of testing the proposed method based on simulation data on the flight test bench of a modern training aircraft for nine different flight modes under conditions of simulating random errors of onboard measurements corresponding to the flight experiment are presented.
6
Nowacki, Marcin, and Damian Olejniczak. "Determination of the operational parameters values for Airbus A300-600ST Beluga aircraft on the basis of CFD tests." MATEC Web of Conferences 357 (2022): 02019. http://dx.doi.org/10.1051/matecconf/202235702019.
Повний текст джерелаАнотація:
The article presents a method for estimating the values of basic operational and aerodynamic parameters of an aircraft. It contains a multistage analysis of CFD test results for the Beluga Airbus A300-600ST model. The first step of the method is to determine the values of aerodynamic parameters such as lift coefficient, drag coefficient, lift force and drag force in specific flight conditions. A comparative analysis of the coefficients of lift and drag force depending on the angle of attack of the aircraft allowed the estimation of the optimal angle of attack of the Airbus A300-600ST Beluga. In the next step, the operational values and the maximum flight ceiling of the aircraft were determined. For this purpose, the results of CFD simulation tests for the optimal angle of attack of the aircraft were used. This article allowed determining the characteristics of the components of aerodynamic force of an airplane depending on the angle of attack, flight altitude and flight speed, determining the optimal angle of attack of the aircraft and calculation of the optimal and maximum flight ceiling values of the Airbus A300-600ST Beluga.
7
Morelli, Eugene A. "Real-Time Aerodynamic Parameter Estimation Without Air Flow Angle Measurements." Journal of Aircraft 49, no. 4 (July 2012): 1064–74. http://dx.doi.org/10.2514/1.c031568.
Повний текст джерела
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Chen, D. H., W. H. Wu, J. J. Wang, and Y. Huang. "Investigation on the aerodynamic performance of an ejection seat." Aeronautical Journal 111, no. 1120 (June 2007): 373–80. http://dx.doi.org/10.1017/s0001924000004620.
Повний текст джерелаАнотація:
Abstract A unique experimental method is used, in combination with numerical calculation and engineering estimation, to study the aerodynamic performance of an ejection seat at M = 0·60, 0·90 and 1·20, angles-of-attack α = 0°~360°, and sideslip angles (β = 0°~–90°. Several basic characteristics of the aerodynamic performance are explored. The normal force of the ejection seat varies in a sinusoidal way and the axial force in a cosinoidal way, with the angle-of-attack. The model is statically unstable longitudinally at most attitude angles and the longitudinal stability could be improved by a stabiliser. These characteristics result from a large low pressure area caused by the leeward separation and the windward high pressure area in the ejection seat flow field, at all α, due to the blunt configuration. A set of engineering calculation formulae is deduced, based on the aerodynamic characteristics of the ejection seat.
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Lerro, Angelo, Manuela Battipede, Piero Gili, and Alberto Brandl. "Aerodynamic angle estimation: comparison between numerical results and operative environment data." CEAS Aeronautical Journal 11, no. 1 (September 4, 2019): 249–62. http://dx.doi.org/10.1007/s13272-019-00417-x.
Повний текст джерела
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Saderla, S., R. Dhayalan, and A. K. Ghosh. "Non-linear aerodynamic modelling of unmanned cropped delta configuration from experimental data." Aeronautical Journal 121, no. 1237 (January 12, 2017): 320–40. http://dx.doi.org/10.1017/aer.2016.124.
Повний текст джерелаАнотація:
ABSTRACTThe paper presents the aerodynamic characterization of a low-speed unmanned aerial vehicle, with cropped delta planform and rectangular cross section, at and around high angles-of-attack using flight test methods. Since the linear models used for identification from flight data at low and moderate angles of attack become unsuitable for accurate parameter estimation at high angles of attack, a non-linear aerodynamic model has to be considered. Therefore, the Kirchhoff's flow separation model was used to incorporate the non-linearity in the aerodynamic model in terms of flow separation point and stall characteristic parameters. The Maximum Likelihood (ML) and Neural Gauss-Newton (NGN) methods were used to perform the parameter estimation on one set of low angle-of-attack and one set of near-stall flight data. It is evident from the estimates that the NGN method, which does not involve solving equations of motion, performs on a par with the classical ML method. This may be attributed to the reason that NGN method uses a neural network which has been trained by performing point to point mapping of the measured flight data. This feature of NGN method enhances its application over a wider envelope of high angles of attack flight data.
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Agrawal, S., D. Gobiha, and N. K. Sinha. "Nonlinear parameter estimation of airship using modular neural network." Aeronautical Journal 124, no. 1273 (October 29, 2019): 409–28. http://dx.doi.org/10.1017/aer.2019.125.
Повний текст джерелаАнотація:
AbstractThe prime focus of this work is to estimate stability and control derivatives of an airship in a completely nonlinear environment. A complete six degrees of freedom airship model has its aerodynamic model as nonlinear functions of angle of attack. Estimating the parameters of aerodynamic model in a nonlinear environment is challenging as it demands an exhaustive dataset that could cover the entire regime of operation of airship. In this work, data generation is achieved by simulating the mathematical model of airship for different trim conditions obtained from continuation analysis. The mathematical model is simulated using predicted parameter values obtained using DATCOM methodology. A modular neural network is then trained using back-propagation and Adam optimisation algorithm for each of the aerodynamic coefficients separately. The estimated nonlinear airship parameters are found to be consistent with the DATCOM parameter values which were used for open-loop simulation. This validates the proposed methodology and could be extended to estimate airship parameters from real flight data.
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Pavlov, Stanislav. "ESTIMATION OF AMOUNT CHANGING OF LOCAL AERODYNAMIC RESISTANCE OF ELEMENTS OF VENTILATION SYSTEM OF MINE AT REVERSING BEHAVIOR OF VENTILATION SYSTEM." Interexpo GEO-Siberia 2, no. 4 (2019): 212–19. http://dx.doi.org/10.33764/2618-981x-2019-2-4-212-219.
Повний текст джерелаАнотація:
In the work study results of influence of typical working junction angle on amount changing of local aerodynamic resistances when air flows in different directions is represented. Using finite-element software, aerodynamic parameters of elements of mine ventilation network is obtained and is compared with analytical calculations Interinfluence local aerodynamic resistances located at ventilation network consecutively in normal and reversing behavior is discovered. It allows to increase accuracy of calculation of mine ventilation network at emergency ventilation operation.
13
Lu, Chen, Rong-Bing Li, Jian-Ye Liu, and Ting-Wan Lei. "Air Data Estimation by Fusing Navigation System and Flight Control System." Journal of Navigation 71, no. 5 (April 30, 2018): 1231–46. http://dx.doi.org/10.1017/s037346331800022x.
Повний текст джерелаАнотація:
A novel synthetic air data estimation method without using air data sensors is presented, and the method only relies on the information from the Navigation System (NS) and Flight Control System (FCS). The aircraft's aerodynamic model is also required to make a connection between the FCS control parameters and the NS measurements. The airspeed, angle of attack and sideslip, angular velocity and wind speed are defined as state vectors, and state equations are established through the aircraft's aerodynamic model and dynamics. Linear velocity and angular velocity provided by the navigation system are considered as the measurement vector. To deal with variable wind fields, a novel Initialised Three-step Extended Kalman Filter (ITEKF), which considers the wind speed as an unknown input, is developed to track the variation of wind speed. Simulation results based on a Generic Hypersonic Vehicle (GHV) model are presented and compared with an existing method. Factors affecting the method's accuracy include the navigation system accuracy and the aerodynamic model error, are also discussed.
14
BUNESCU, Ionut, Sterian DANAILA, Mihai-Victor PRICOP, and Adrian DINA. "Estimation of Wind Tunnel Corrections Using Potential Models." INCAS BULLETIN 11, no. 1 (March 5, 2019): 53–60. http://dx.doi.org/10.13111/2066-8201.2019.11.1.4.
Повний текст джерелаАнотація:
The evaluation of the tunnel correction remains an actual problem, especially for the effect of tunnel walls. Even if the experimental campaign meets the basic similitude criteria (Mach, Reynolds etc.), the wall effect on the measured data is always present. Consequently, the flow correction due the limited by walls must be evaluated. Solid wall corrections refer to the aerodynamic interference between the experimental model and the walls of the wind tunnel. This interaction affects the measured forces and implicitly the angle of attack. Usually, these effects are introduced through semi-empirical correction factors which change the global measured forces. The present paper refers to the mathematical and numerical modeling of aerodynamic interferences between the experimental model and the solid walls based on the potential flow model. The main goal is to asses a method allowing an estimate of the corrections for each configuration with a minimum computational resource.
15
Korsun, O. N., S. Yu Prihodko, and S. A. Sergeev. "Estimation of the flight thrust increment changing the operation mode of the engine." ITM Web of Conferences 18 (2018): 01002. http://dx.doi.org/10.1051/itmconf/20181801002.
Повний текст джерелаАнотація:
The paper presents a new method for flight test identification of the thrust increment when changing engine operation mode. It is known that the problem of separate estimation of the engine thrust and the aerodynamic drag force from the aircraft flight data belongs to a class of ill-posed. The idea of the presented method is to come out of the ill-posed class by changing the original problem formulation. To achieve this purpose we give up the absolute values of the thrust and proceed to estimating the thrust increment resulting from the change of the aircraft engine operation mode. The accuracy of the solution is improved by introducing the special flight test maneuver. The algorithm requires only a few on-board measured parameters, such as longitudinal and normal overload, airspeed and altitude, and angle of attack. The complimentary use of the engine gas-dynamic model may improve the estimates but is not obligatory. The proposed method also does not require a-priory estimates of the aerodynamic drag force.
16
Meijer, M. C., L. Dala, and L. Dala. "Quantifying non-linearity in planar supersonic potential flows." Aeronautical Journal 121, no. 1237 (January 18, 2017): 372–94. http://dx.doi.org/10.1017/aer.2016.141.
Повний текст джерелаАнотація:
ABSTRACTAn analysis is presented which allows the engineer to quantitatively estimate the validity bounds of aerodynamic methods based in linear potential flows a-priori. The development is limited to quasi-steady planar flows with attached shocks and small body curvature. Perturbation velocities are parameterised in terms of Mach number and flow turning angle by means of a series-expansion for flow velocity based in the method of characteristics. The parameterisation is used to assess the magnitude of non-linear term-groupings relative to linear groups in the full potential equation. This quantification is used to identify dominant nonlinear terms and to estimate the validity of linearising the potential flow equation at a given Mach number and flow turning angle. Example applications include the a-priori estimation of the validity bounds for linear aerodynamic models for supersonic aeroelastic analysis of lifting surfaces and panels.
17
Xia, X., and K. Mohseni. "Unsteady aerodynamics and vortex-sheet formation of a two-dimensional airfoil." Journal of Fluid Mechanics 830 (October 2, 2017): 439–78. http://dx.doi.org/10.1017/jfm.2017.513.
Повний текст джерелаАнотація:
Unsteady inviscid flow models of wings and airfoils have been developed to study the aerodynamics of natural and man-made flyers. Vortex methods have been extensively applied to reduce the dimensionality of these aerodynamic models, based on the proper estimation of the strength and distribution of the vortices in the wake. In such modelling approaches, one of the most fundamental questions is how the vortex sheets are generated and released from sharp edges. To determine the formation of the trailing-edge vortex sheet, the classical steady Kutta condition can be extended to unsteady situations by realizing that a flow cannot turn abruptly around a sharp edge. This condition can be readily applied to a flat plate or an airfoil with cusped trailing edge since the direction of the forming vortex sheet is known to be tangential to the trailing edge. However, for a finite-angle trailing edge, or in the case of flow separation away from a sharp corner, the direction of the forming vortex sheet is ambiguous. To remove any ad hoc implementation, the unsteady Kutta condition, the conservation of circulation as well as the conservation laws of mass and momentum are coupled to analytically solve for the angle, strength and relative velocity of the trailing-edge vortex sheet. The two-dimensional aerodynamic model together with the proposed vortex-sheet formation condition is verified by comparing flow structures and force calculations with experimental results for several airfoil motions in steady and unsteady background flows.
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Zhong, Wei, Wen Zhong Shen, Tong Guang Wang, and Wei Jun Zhu. "A New Method of Determination of the Angle of Attack on Rotating Wind Turbine Blades." Energies 12, no. 20 (October 22, 2019): 4012. http://dx.doi.org/10.3390/en12204012.
Повний текст джерелаАнотація:
The angle of attack (AoA) is the key parameter when extracting the aerodynamic polar from the rotating blade sections of a wind turbine. However, the determination of AoA is not straightforward using computational fluid dynamics (CFD) or measurement. Since the incoming streamlines are bent because of the complex inductions of the rotor, discrepancies exist between various existing determination methods, especially in the tip region. In the present study, flow characteristics in the region near wind turbine blades are analyzed in detail using CFD results of flows past the NREL UAE Phase VI rotor. It is found that the local flow determining AOA changes rapidly in the vicinity of the blade. Based on this finding, the concepts of effective AoA as well as nominal AoA are introduced, leading to a new method of AOA determination. The new method has 5 steps: (1) Find the distributed vortices on the blade surface; (2) select two monitoring points per cross-section close to the aerodynamic center on both pressure and suction sides with an equal distance from the rotor plane; (3) subtract the blade self-induction from the velocity at each monitoring point; (4) average the velocity of the two monitoring points obtained in Step 3; (5) determine the AoA using the velocity obtained in Step 4. Since the monitoring points for the first time can be set very close to the aerodynamic center, leading to an excellent estimation of AoA. The aerodynamic polar extracted through determination of the effective AoA exhibits a consistent regularity for both the mid-board and tip sections, which has never been obtained by the existing determination methods.
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Lerro, Angelo, Alberto Brandl, Manuela Battipede, and Piero Gili. "Preliminary Design of a Model-Free Synthetic Sensor for Aerodynamic Angle Estimation for Commercial Aviation." Sensors 19, no. 23 (November 23, 2019): 5133. http://dx.doi.org/10.3390/s19235133.
Повний текст джерелаАнотація:
Heterogeneity of the small aircraft category (e.g., small air transport (SAT), urban air mobility (UAM), unmanned aircraft system (UAS)), modern avionic solution (e.g., fly-by-wire (FBW)) and reduced aircraft (A/C) size require more compact, integrated, digital and modular air data system (ADS) able to measure data from the external environment. The MIDAS project, funded in the frame of the Clean Sky 2 program, aims to satisfy those recent requirements with an ADS certified for commercial applications. The main pillar lays on a smart fusion between COTS solutions and analytical sensors (patented technology) for the identification of the aerodynamic angles. The identification involves both flight dynamic relationships and data-driven state observer(s) based on neural techniques, which are deterministic once the training is completed. As this project will bring analytical sensors on board of civil aircraft as part of a redundant system for the very first time, design activities documented in this work have a particular focus on airworthiness certification aspects. At this maturity level, simulated data are used, real flight test data will be used in the next stages. Data collection is described both for the training and test aspects. Training maneuvers are defined aiming to excite all dynamic modes, whereas test maneuvers are collected aiming to validate results independently from the training set and all autopilot configurations. Results demonstrate that an alternate solution is possible enabling significant savings in terms of computational effort and lines of codes but they show, at the same time, that a better training strategy may be beneficial to cope with the new neural network architecture.
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Leong Chee Hang, Nur Athirah Nadwa Rosli, Mastura Ab Wahid, Norazila Othman, Shabudin Mat, and Mohd Zarhamdy Md Zain. "CFD Analysis on Propeller at Varying Propeller Disc Angle and Advance Ratio." Journal of Advanced Research in Fluid Mechanics and Thermal Sciences 96, no. 1 (July 6, 2022): 82–95. http://dx.doi.org/10.37934/arfmts.96.1.8295.
Повний текст джерелаАнотація:
The purpose of this study is to see the feasibility of using Computational Fluid Dynamic (CFD) analysis over wind tunnel testing for propeller performance measurement. Computational Fluid Dynamic (CFD) analysis is conducted on APC 6X4E propeller and 9X6E propeller at propeller disc angle, α of 0°, 30°, 60°, and 90° using commercially available software ANSYS FLUENT at advance ratio ranging from 0 to 0.88 and angular velocity ranging from 1000rpm to 8000rpm. CFD analysis was also performed for different advance ratio at different rotational speed of 4000rpm and 8000rpm. Multiple Reference Frame model is used to simulate the rotating propeller in a flowing airstream by constructing a rotating and static domain around the propeller. The non-zero propeller disc angle is achieved by changing the inlet direction of the static domain. The SST k-ω turbulence model is used, and tetrahedral mesh is constructed. The propeller thrust and torque obtained from the CFD simulation is used to calculate the aerodynamic characteristics of the propellers. The thrust coefficient, torque coefficient and propeller efficiency obtained for both propellers follow the trend of the wind tunnel testing. The results obtained from the CFD simulation matches with the results trend obtained by another researcher performing wind tunnel analysis, where the thrust coefficient decreases with increasing advance ratio at propeller disc angle less than 60° and increases with increasing advance ratio at propeller disc angle more than 70°. The error produce for thrust estimation is lower than 12% for both 6x4E and 9x6E propellers. The error for torque and efficiency estimation is between 15-30% and 12% respectively. In conclusion, CFD simulation can predict the aerodynamic characteristics of Low Reynolds Number propeller at different propeller disc angle.
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Zhong, Siping, and Wenfu Xu. "Power Modeling and Experiment Study of Large Flapping-Wing Flying Robot during Forward Flight." Applied Sciences 12, no. 6 (March 21, 2022): 3176. http://dx.doi.org/10.3390/app12063176.
Повний текст джерелаАнотація:
A power estimation approach for calculating the power of a flapping-wing air vehicle (FWAV) in forward flight is proposed in this paper. One of the challenges and essential points of FWAVs is endurance. In order to optimize FWAVs, it is necessary to analyze power required for flight in addition to kinematic and aerodynamic analyses of the prototype. Previously, calculating the power of birds was limited to calculating their average power, which assumed the lift was usually the same as the gravitational force. However, the lift varies with the flapping angle during flight. As a result, the power required for forward flight of FWAVs is determined in this work by using a kinematic model of the drive element and wing flapping, along with the aerodynamic model, which varies with the flapping angle during the flapping cycle. Experiments were performed with two prototypes with wingspans of 1.6 and 1.8 m, utilizing a wind tunnel platform. The correlations between power and angle of attack, flapping frequency, and incoming flow velocity were discovered, and recommendations for FWAVs and flying mode design were provided. However, several challenges are highlighted in the application of the model to practical design efforts.
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Gao, Zhenxing, Haofeng Wang, Zhiwei Xiang, and Debao Wang. "Flight Data-Based Wind Disturbance and Air Data Estimation." Atmosphere 12, no. 4 (April 8, 2021): 470. http://dx.doi.org/10.3390/atmos12040470.
Повний текст джерелаАнотація:
The instantaneous wind field and air data, including true airspeed, angle of attack, angle of sideslip, cannot be measured and recorded accurately in wind disturbance. A new air data and wind field estimation method is proposed based on flight data in this study. Since the wind field is the horizontal prevailing wind added by turbulence, the slowly time-varying prevailing wind and small-scale turbulence are described by the exponentially correlated stochastic wind model and von Karman turbulence model, respectively. The system update equation of air data is built based on inertial measurements instead of the complex aerodynamic and aero-engine model of aircraft. Benefitted by the post-analysis characteristics of flight data, a forward–backward filtering algorithm was designed to improve the estimation accuracy. Simulation results indicate that the forward–backward filter integrated with the von Karman turbulence model can reduce the estimation error and ensure filtering stability. A further test with actual flight data shows that the forward–backward filter is not only able to track the wide-range change in prevailing wind but also reduce the adverse effects of uncertain disturbance on estimation accuracy.
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Billingsley, Ethan, Mehdi Ghommem, Rui Vasconcellos, and Abdessattar Abdelkefi. "On the Aerodynamic Analysis and Conceptual Design of Bioinspired Multi-Flapping-Wing Drones." Drones 5, no. 3 (July 18, 2021): 64. http://dx.doi.org/10.3390/drones5030064.
Повний текст джерелаАнотація:
Many research studies have investigated the characteristics of bird flights as a source of bioinspiration for the design of flapping-wing micro air vehicles. However, to the best of the authors’ knowledge, no drone design targeted the exploitation of the aerodynamic benefits associated with avian group formation flight. Therefore, in this work, a conceptual design of a novel multi-flapping-wing drone that incorporates multiple pairs of wings arranged in a V-shape is proposed in order to simultaneously increase the propulsive efficiency and achieve superior performance. First, a mission plan is established, and a weight estimation is conducted for both 3-member and 5-member configurations of the proposed air vehicle. Several wing shapes and airfoils are considered, and aerodynamic simulations are conducted, to determine the optimal planform, airfoil, formation angle, and angle of attack. The simulation results reveal that the proposed bioinspired design can achieve a propulsive efficiency of 73.8%. A stability analysis and tail sizing procedure are performed for both 3-member and 5-member configurations. In addition, multiple flapping mechanisms are inspected for implementation in the proposed designs. Finally, the completed prototypes’ models of the proposed multi-flapping-wing air vehicles are presented, and their features are discussed. The aim of this research is to provide a framework for the conceptual design of bioinspired multi-flapping-wing drones and to demonstrate the sizing, weight estimation, and design procedures for this new type of air vehicles. This work establishes the first multi-flapping-wing drone design which exploits the aerodynamic features of the V-formation flight observed in birds to achieve superior performance in terms of payload and endurance.
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Valldosera Martinez, Robert, Frederico Afonso, and Fernando Lau. "Aerodynamic Shape Optimisation of a Camber Morphing Airfoil and Noise Estimation." Aerospace 9, no. 1 (January 15, 2022): 43. http://dx.doi.org/10.3390/aerospace9010043.
Повний текст джерелаАнотація:
In order to decrease the emitted airframe noise by a two-dimensional high-lift configuration during take-off and landing performance, a morphing airfoil has been designed through a shape design optimisation procedure starting from a baseline airfoil (NLR 7301), with the aim of emulating a high-lift configuration in terms of aerodynamic performance. A methodology has been implemented to accomplish such aerodynamic improvements by means of the compressible steady RANS equations at a certain angle of attack, with the objective of maximising its lift coefficient up to equivalent values regarding the high-lift configuration, whilst respecting the imposed structural constraints to guarantee a realistic optimised design. For such purposes, a gradient-based optimisation through the discrete adjoint method has been undertaken. Once the optimised airfoil is achieved, unsteady simulations have been carried out to obtain surface pressure distributions along a certain time-span to later serve as the input data for the aeroacoustic prediction framework, based on the Farassat 1A formulation, where the subsequent results for both configurations are post-processed to allow for a comparative analysis. Conclusively, the morphing airfoil has proven to be advantageous in terms of aeroacoustics, in which the noise has been reduced with respect to the conventional high-lift configuration for a comparable lift coefficient, despite being hampered by a significant drag coefficient increase due to stall on the morphing airfoil’s trailing edge.
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Moriche, M., M. Raiola, S. Discetti, A. Ianiro, O. Flores, and M. García-Villalba. "Assessing aerodynamic force estimation with experiments and simulations of flapping-airfoil flows on the verge of three-dimensionality." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 234, no. 2 (August 19, 2019): 428–44. http://dx.doi.org/10.1177/0954410019867570.
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This paper reports a combined experimental and numerical study of the flow over a rigid airfoil in flapping motion. The setup consists of a heaving and pitching airfoil at a moderate Reynolds number ([Formula: see text]), at a Strouhal number St = 0.1. The aim is to assess the accuracy of two-dimensional direct numerical simulations in predicting aerodynamic forces in a flow configuration, which is nominally two-dimensional but is at the verge of three-dimensionality. The assessment is carried out with experiments, including flow field and aerodynamic force measurements with particle image velocimetry and a load cell. The comparative study shows a good qualitative agreement between the experiments and the simulations at comparable Reynolds numbers both in terms of forces and flow fields, but with some quantitative differences. The quantitative discrepancies between experiments and simulation are analyzed and reduced to inherent differences between experimental and computational setups. It is observed that the significant differences are apparent almost exclusively in the wake evolution. Nonetheless, this is shown to have a minor effect on the aerodynamic force estimation. Overall, the trends observed when varying the mean pitch angle and the pitching amplitude are the same in both experiments and simulations. This suggests that two-dimensional/three-dimensional effects do not alter significantly the relationship between the unsteady flow mechanisms (i.e. leading edge vortex) and the aerodynamic forces in the parametric range considered here.
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Han, Jong-Seob, and Christian Breitsamter. "Aerodynamic investigation on shifted-back vertical stroke plane of flapping wing in forward flight." Bioinspiration & Biomimetics 16, no. 6 (November 1, 2021): 064001. http://dx.doi.org/10.1088/1748-3190/ac305f.
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Abstract In order to properly understand aerodynamic characteristics in a flapping wing in forward flight, additional aerodynamic parameters apart from those in hover—an inclined stroke plane, a shifted-back stroke plane, and an advance ratio—must be comprehended in advance. This paper deals with the aerodynamic characteristics of a flapping wing in a shifted-back vertical stroke plane in freestream. A scaled-up robotic arm in a water towing tank was used to collect time-varying forces of a model flapping wing, and a semi-empirical quasi-steady aerodynamic model, which can decompose the forces into steady, quasi-steady, and unsteady components, was used to estimate the forces of the model flapping wing. It was found that the shifted-back stroke plane left a part of freestream as a non-perpendicular component, giving rise to a time-course change in the aerodynamic forces during the stroke. This also brought out two quasi-steady components (rotational and added-mass forces) apart from the steady one, even the wing moved with a constant stroke velocity. The aerodynamic model underestimated the actual forces of the model flapping wing even it can cover the increasingly distributed angle of attack of the vertical stroke plane with a blade element theory. The locations of the centers of pressure suggested a greater pressure gradient and an elongated leading-edge vortex along a wingspan than that of the estimation, which may explain the higher actual force in forward flight.
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Yang, Wenqing, Bifeng Song, and Guanglin Gao. "Flight Performance Estimation of Bionic Flapping-Wing Micro Air Vehicle." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 36, no. 4 (August 2018): 636–43. http://dx.doi.org/10.1051/jnwpu/20183640636.
Повний текст джерелаАнотація:
Bionic flapping-wing micro air vehicle(MAV) has received worldwide attention.The flight performance calculation is an important step in the conceptual design.The differences in performance estimation methods between the flapping-wing and conventional fixed-wing aircraft are analyzed.Based on the results of the aerodynamic estimation and wind tunnel experimental measurement, the flight performance estimation method of flapping-wing micro air vehicle is proposed, and the performance of level flight, climbing, and duration are calculated and analyzed.The frequency represents the accelerator in a certain extent, while the frequency is coupled with lift and thrust.The results show that there may be two stable cruising states at certain frequencies, one is the small angle of attack with high speed, the other is the small speed with big angle of attack, and the two states have different power consumption.According to the parameters of the vehicle, climbing performance and duration performance can be obtained.The speed versus power characteristic curve is a U shape, minimum slope of the U curve can be obtained through the mapping method to calculate the farthest flight speed, and the minimum velocity of U-shaped curve is the speed for longest duration.The proposed flight performance calculation method can be used to evaluate the flight capability of bionic micro flapping-wing air vehicle.
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Duport, Chloé, Jean-Baptiste Leroux, Kostia Roncin, Christian Jochum, and Yves Parlier. "Benchmarking of a 3D non-linear lifting line method against 3D RANSE simulations." La Houille Blanche, no. 5-6 (December 2019): 70–73. http://dx.doi.org/10.1051/lhb/2019029.
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As a part of the design and operation of kites as auxiliary propulsion of vessels, it is necessary to be able to quickly estimate the aerodynamic efforts along various trajectories. A 3D non-linear model based on the lifting line of Prandtl has been developed for this purpose. It allows these rapid calculations for wings with any laws for the dihedral angle, the twist, and the sweep angle, along the span, and for a general flight kinematic taking into account translation velocities and rotation rates. This model has been verified by comparison with 3D simulations performed with a Navier-Stokes solver. It gives satisfactory results in incidence and sideslip, with gaps of about 4% for forecasts lift. Special attention has been paid to the estimation of the accuracy of the provided numerical results.
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Afanasieva, Nadiia. "The Effect of Angle of Attack and Flow Conditions on Turbulent Boundary Layer Noise of Small Wind Turbines." Archives of Acoustics 42, no. 1 (March 1, 2017): 83–91. http://dx.doi.org/10.1515/aoa-2017-0009.
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Abstract The article aims to solve the problem of noise optimization of small wind turbines. The detailed analysis concentrates on accurate specification and prediction of the turbulent boundary layer noise spectrum of the blade airfoil. The angles of attack prediction for a horizontal axis wind turbine (HAWT) and the estimation based on literature data for a vertical axis one (VAWT), were conducted, and the influence on the noise spectrum was considered. The 1/3-octave sound pressure levels are obtained by semi-empirical model BPM. Resulting contour plots show a fundamental difference in the spectrum of HAWT and VAWT reflecting the two aerodynamic modes of flow that predefine the airfoil self-noise. Comparing the blade elements with a local radius of 0.875 m in the HAWT and VAWT conditions the predicted sound pressure levels are the 78.5 dB and 89.8 dB respectively. In case of the HAWT with predicted local angle of attack ranging from 2.98° to 4.63°, the acoustic spectrum will vary primarily within broadband frequency band 1.74-20 kHz. For the VAWT with the local angle of attack ranging from 4° to 20° the acoustic spectrum varies within low and broadband frequency bands 2 Hz - 20 kHz.
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Bahrami, M., B. Ebrahimi, and G. R. Ansarifar. "Sliding Mode Observer and Control Design with Adaptive Parameter Estimation for a Supersonic Flight Vehicle." International Journal of Aerospace Engineering 2010 (2010): 1–9. http://dx.doi.org/10.1155/2010/474537.
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Design and synthesis of a nonlinear generic supersonic flight vehicle longitudinal dynamics control for angle-of-attack, AOA, output tracking in the atmospheric flight is presented based on sliding mode control. A sliding mode observer is invoked to estimate AOA which is difficult to measure in practice. Large parameter uncertainties accommodation envisaged by designing adaptive mechanisms for both the control and observer and high chattering authority due to large deviations of aerodynamic coefficients arising from wind-tunnel measurements are inhibited. The employed method enables the sliding mode control design to exhibit the desired dynamic properties during the entire output-tracking process. Simulations results are presented to demonstrate the performance, robustness, and stability.
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Mehta, R. C. "Estimation of Angle of Attack in Satellite Launch Vehicle Using Flush Air Data Sensing Systems at Mach 0.5 to 3.0." Scholars Journal of Engineering and Technology 9, no. 7 (August 1, 2021): 77–86. http://dx.doi.org/10.36347/sjet.2021.v09i07.001.
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This paper presents an inverse analysis to estimate angle of attack during ascent period of a satellite launch vehicle. Aerodynamic results are numerically computed by solving three dimensional, time dependent, compressible inviscid equations over payload shroud of a satellite launch vehicle. The flush air data system consists of four pressure ports flushed with conical-nose section of the payload fairing and connected to on board differential pressure transducers. The inverse algorithm uses calibrations charts which based on computed and measured data. A controlled random search method is used to predict pitch, yaw and total angle of attack of vehicle from measured transient differential pressure history in flight from Mach numbers range of 0.5 to 3.0. The algorithm predicts the flow direction stepwise with function of flight Mach numbers and can be termed as online method. Flow direction of the launch vehicle is compared with the reconstructed trajectory data. The estimated values of the flow direction are found in good agreement them.
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Pontillo, Alessandro, Sezsy Yusuf, Guillermo Lopez, Dominic Rennie, and Mudassir Lone. "Investigating pitching moment stall through dynamic wind tunnel test." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 234, no. 2 (July 19, 2019): 267–79. http://dx.doi.org/10.1177/0954410019861853.
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Experimental characterisation of aircraft dynamic stall can be a challenging and complex system identification activity. In this article, the authors present a method that combines dynamic wind tunnel testing with parameter estimation techniques to study the nonlinear pitching moment dynamics of a 1/12 scale Hawk model undergoing moment stall. The instrumentation setup allows direct calculation of angular acceleration terms, such as pitch acceleration, and avoids post-processing steps involving differentiation of signals. Data collected from tests, carried out at 20 m/s and 30 m/s, are used for a brief aerodynamic analysis of the observed stall hysteresis. Then an output-error-based parameter estimation process is used to parameterise dynamic stall models and furthermore, illustrate that in a scenario where the model's heave motion is constrained. The observed nonlinear behaviour arises from the nonlinear angle of attack and linear pitch rate components.
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Gros, Sébastien, Hammad Ahmad, Kurt Geebelen, Jan Swevers, and Moritz Diehl. "In-flight Estimation of the Aerodynamic Roll Damping and Trim Angle for a Tethered Aircraft based on Multiple-shooting*." IFAC Proceedings Volumes 45, no. 16 (July 2012): 1407–12. http://dx.doi.org/10.3182/20120711-3-be-2027.00342.
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Kshitij, A., S. A. Prince, J. L. Stollery, and F. de la P. Ricón. "A simple method for drag estimation for wedge-like fairings in hypersonic flow." Aeronautical Journal 125, no. 1288 (March 23, 2021): 968–87. http://dx.doi.org/10.1017/aer.2021.20.
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ABSTRACTThe addition of wedge-like fairings onto the side of missiles and space launch vehicles, to shield devices such as cameras and reaction jet nozzles, creates additional drag, particularly when in supersonic and hypersonic freestream flow. An experimental and computational study was performed to obtain aerodynamic data on simple representative configurations to test the accuracy of simple theories for the drag increment due to these types of fairings. A semi-empirical method to estimate drag on wedge-shaped projections is presented, which may be used by missile designers to provide predictions of the drag increment due to wedge-like fairings. The method is shown to be valid where the wedge width is much smaller than body diameter, and across the Mach number range 4–8.2 but is likely to be valid for higher Mach numbers. Drag coefficient is found to increase with increasing wedge angle and reducing wedge slenderness, although increasing slenderness tends to increase skin friction drag.
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MATSUOKA, Yasuaki, and Koichi YONEMOTO. "E-06 Estimation Accuracy Evaluation on Aerodynamic Characteristics of Three-dimensional Wing by Wake Integration Method for High-Angle-Flow." Proceedings of Conference of Kyushu Branch 2016.69 (2016): 183–84. http://dx.doi.org/10.1299/jsmekyushu.2016.69.183.
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Liu, Xu, Wei Liu, and Yun Fei Zhao. "Research Progress and Prospects for Vehicle Dynamic Stability Parameters." Applied Mechanics and Materials 729 (January 2015): 95–100. http://dx.doi.org/10.4028/www.scientific.net/amm.729.95.
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Dynamic stability parameters (dynamic derivatives) are important indicators for the control system design, orbit design and longitudinal and horizontal dynamic stability analysis of aircrafts. Methods that evaluate the quality and dynamics of an aircraft typically include flight experiment, wind tunnel testing and theoretical calculation, with one of the most important part of them being the obtainment of dynamic derivatives. Project estimation method derivative action is not considered suitable for boundary layer transition, flow separation and re-attached and the complex situation leeward area vortex small angle of attack linear range. Frequency domain is a dynamic non-scheduled periodic invariant system to get moving derivative calculation method, but the accuracy of the unsteady flow is much lower than the time-domain calculations. Currently, unsteady CFD approach represents a time-domain nonlinear aerodynamic characteristics predicted the most advanced level. Derivative prediction efficiency and adaptability under conditions of high angle of attack of the development trend of nonlinear dynamic derivatives were analyzed. As a global trend, obtaining dynamic parameters through numerical calculation is becoming a prevailing approach to dynamic parameter research.
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Shi, Guoxiang, Ke Zhang, Pei Wang, and Zhiguo Han. "Algorithm of Reentry Guidance for Hypersonic Vehicle Based on Lateral Maneuverability Prediction." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 38, no. 3 (June 2020): 523–32. http://dx.doi.org/10.1051/jnwpu/20203830523.
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Aiming at the problem that the traditional error corridor guidance method has poor adaptability in lateral guidance of predictor-corrector guidance, an algorithm of reentry guidance based on the vehicle lateral maneuverability prediction is proposed without increasing the calculation too much. The lateral component mean value of lift at reentry is calculated by using the bank angle magnitude function obtained from longitudinal guidance. According to the above-mentioned, a crossrange corridor with dynamic boundary constraint is designed to control bank angle reversal timing. Online parameters estimation is introduced to suppress the influence of the atmospheric density and aerodynamic parameters disturbance on the predictor model. The CAV-L, a kind of hypersonic vehicle, is used as an object to carry out reentry guidance simulation. The results show that the guidance algorithm can effectively guide vehicle to target for reentry missions of different range, the landing point error are small and the guidance effect is stable. The simulated results via Monte Carlo method verify that the guidance algorithm has a good adaptability and robustness to initial state deviations and process disturbances.
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Ahmad, Muhammad, Zukhruf Liaqat Hussain, Syed Irtiza Ali Shah, and Taimur Ali Shams. "Estimation of Stability Parameters for Wide Body Aircraft Using Computational Techniques." Applied Sciences 11, no. 5 (February 26, 2021): 2087. http://dx.doi.org/10.3390/app11052087.
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In this paper, we present the procedure of estimating the aerodynamic coefficients for a commercial aviation aircraft from geometric parameters at low-cruise-flight conditions using US DATCOM (United States Data Compendium) and XFLR software. The purpose of this research was to compare the stability parameters from both pieces of software to determine the efficacy of software solution for a wide-body aircraft at the stated flight conditions. During the initial phase of this project, the geometric parameters were acquired from established literature. In the next phase, stability and control coefficients of the aircraft were estimated using both pieces of software in parallel. Results obtained from both pieces of software were compared for any differences and the both pieces of software were validated with analytical correlations as presented in literature. The plots of various parameters with variations of the angle of attack or control surface deflection have also been obtained and presented. The differences between the software solutions and the analytical results can be associated with approximations of techniques used in software (the vortex lattice method is the background theory used in both DATCOM and XFLR). Additionally, from the results, it can be concluded that XFLR is more reliable than DATCOM for longitudinal, directional, and lateral stability/control coefficients. Analyses of a Boeing 747-200 (a wide-body commercial airliner) in DATCOM and XFLR for complete stability/control analysis including all modes in the longitudinal and lateral directions have been presented. DATCOM already has a sample analysis of a previous version of the Boeing 737; however, the Boeing 747-200 is much larger than the former, and complete analysis was, therefore, felt necessary to study its aerodynamics characteristics. Furthermore, in this research, it was concluded that XFLR is more reliable for various categories of aircraft alike in terms of general stability and control coefficients, and hence many aircraft can be dependably modeled and analyzed in this software.
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Juliawan, Nadhie, Hyoung-Seog Chung, Jae-Woo Lee, and Sangho Kim. "Estimation and Separation of Longitudinal Dynamic Stability Derivatives with Forced Oscillation Method Using Computational Fluid Dynamics." Aerospace 8, no. 11 (November 19, 2021): 354. http://dx.doi.org/10.3390/aerospace8110354.
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This paper focuses on estimating dynamic stability derivatives using a computational fluid dynamics (CFD)-based force oscillation method, and on separating the coupled dynamic derivatives terms obtained from the method. A transient RANS solver is used to calculate the time history of aerodynamic moments for a test model oscillating about the center of gravity, from which the coupled dynamic derivatives are estimated. The separation of the coupled derivatives term is carried out by simulating simple harmonic oscillation motions such as plunging motion and flapping motion which can isolate the pitching moment due to AOA rate (Cmα˙) and the pitching moment due to pitch rate (Cmq), respectively. The periodic motions are implemented using a CFD dynamic mesh technique with user-defined function (UDF). For the validation test, steady and unsteady simulations are performed on the Army-Navy Finner Missile model. The static aerodynamic moments and pressure distribution, as well as the coupled dynamic derivative results from the pitching oscillation mode, show good agreement with the previously published wind tunnel tests and CFD analysis data. In order to separate the coupled derivative terms, two additional harmonic oscillation modes of plunging and flapping motions are tested with the angle of attack variations from 0 to 85 degrees at a supersonic speed to provide real insight on the missile maneuverability. The cross-validation study between the three oscillation modes indicates the summation of the individual plunging and flapping results becoming nearly identical to the coupled derivative results from the pitching motion, which implies the entire set of coupled and separated dynamic derivative terms can be effectively estimated with only two out of three modes. The advantages and disadvantages of each method are discussed to determine the efficient approach of estimating the dynamic stability derivatives using the forced oscillation method.
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Corcione, Salvatore, Vincenzo Cusati, Danilo Ciliberti, and Fabrizio Nicolosi. "Experimental Assessment of Aero-Propulsive Effects on a Large Turboprop Aircraft with Rear-Engine Installation." Aerospace 10, no. 1 (January 15, 2023): 85. http://dx.doi.org/10.3390/aerospace10010085.
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This paper deals with the estimation of propulsive effects for a three-lifting surface turboprop aircraft concept, with rear engine installation at the horizontal tail tips, conceived to carry up to 130 passengers. This work is focused on how the propulsive system affects the horizontal tailplane aerodynamics and, consequently, the aircraft’s static stability characteristics using wind tunnel tests. Both direct and indirect propulsive effects have been estimated. The former produces moments whose values depend on the distance from the aircraft’s centre of gravity to the thrust lines and propeller disks. The latter entails a change in the angle of attack and an increment of dynamic pressure on the tailplane. Several tests were also performed on the body-empennage configuration to investigate the propulsive effects on the aircraft’s static stability without the appearance of any aerodynamic interference phenomena, especially from the canard. The output of the experimental campaign reveals a beneficial effect of the propulsive effects on the aircraft’s longitudinal stability, with an increase in the stability margin of about 2.5% and a reduction in the directional stability derivative of about 4%, attributed to the different induced drag contributions of the two horizontal tail semi-planes. Moreover, the rolling moment coefficient experiences a greater variation due to the propulsion depending on the propeller rotation direction. The outcomes of this paper allow the enhancement of the technical readiness level for the considered aircraft, giving clear indications about the feasibility of the aircraft configuration.
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Thedens, Paul, and Roland Schmehl. "An Aero-Structural Model for Ram-Air Kite Simulations." Energies 16, no. 6 (March 9, 2023): 2603. http://dx.doi.org/10.3390/en16062603.
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Similar to parafoils, ram-air kites are flexible membrane wings inflated by the apparent wind and supported by a bridle line system. A major challenge in estimating the performance of these wings using a computer model is the strong coupling between the airflow around the wing and the deformation of the membrane structure. In this paper, we introduce a staggered coupling scheme combining a structural finite element solver using a dynamic relaxation technique with a potential flow solver. The developed method proved numerically stable for determining the equilibrium shape of the wing under aerodynamic load and is thus suitable for performance measurement and load estimation. The method was validated with flight data provided by SkySails Power. Measured forces on the tether and steering belt of the robotic kite control pod showed good resemblance with the simulation results. As expected for a potential flow solver, the kite’s glide ratio was overestimated by 10–15%, and the measured tether elevation angle in a neutral flight scenario matched the simulations within 2 degrees. Based on the obtained results, it can be concluded that the proposed aero-structural model can be used for initial designs of ram-air kites with application to airborne wind energy.
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Abdessemed, Chawki, Abdessalem Bouferrouk, and Yufeng Yao. "Aerodynamic and Aeroacoustic Analysis of a Harmonically Morphing Airfoil Using Dynamic Meshing." Acoustics 3, no. 1 (March 6, 2021): 177–99. http://dx.doi.org/10.3390/acoustics3010013.
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This work explores the aerodynamic and aeroacoustic responses of an airfoil fitted with a harmonically morphing Trailing Edge Flap (TEF). An unsteady parametrization method adapted for harmonic morphing is introduced, and then coupled with dynamic meshing to drive the morphing process. The turbulence characteristics are calculated using the hybrid Stress Blended Eddy Simulation (SBES) RANS-LES model. The far-field tonal noise is predicted using the Ffowcs-Williams and Hawkings (FW-H) acoustic analogy method with corrections to account for spanwise effects using a correlation length of half the airfoil chord. At various morphing frequencies and amplitudes, the 2D aeroacoustic tonal noise spectra are obtained for a NACA 0012 airfoil at a low angle of attack (AoA = 4°), a Reynolds number of 0.62 × 106, and a Mach number of 0.115, respectively, and the dominant tonal frequencies are predicted correctly. The aerodynamic coefficients of the un-morphed configuration show good agreement with published experimental and 3D LES data. For the harmonically morphing TEF case, results show that it is possible to achieve up to a 3% increase in aerodynamic efficiency (L/D). Furthermore, the morphing slightly shifts the predominant tonal peak to higher frequencies, possibly due to the morphing TEF causing a breakup of large-scale shed vortices into smaller, higher frequency turbulent eddies. It appears that larger morphing amplitudes induce higher sound pressure levels (SPLs), and that all the morphing cases induce the shift in the main tonal peak to a higher frequency, with a maximum 1.5 dB reduction in predicted SPL. The proposed dynamic meshing approach incorporating an SBES model provides a reasonable estimation of the NACA 0012 far-field tonal noise at an affordable computational cost. Thus, it can be used as an efficient numerical tool to predict the emitted far-field tonal noise from a morphing wing at the design stage.
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Zhang, Huang, Liu, and Zhang. "Fuel Consumption Model of the Climbing Phase of Departure Aircraft Based on Flight Data Analysis." Sustainability 11, no. 16 (August 12, 2019): 4362. http://dx.doi.org/10.3390/su11164362.
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Accurate estimation of the fuel consumed during aircraft operation is key for determining the fuel load, reducing the airline operating cost, and mitigating environmental impacts. Aerodynamic parameters in current fuel consumption models are obtained from a static diagram extracted from the outcomes of wind tunnel experiments. Given that these experiments are performed in a lab setting, the parameters cannot be used to estimate additional fuel consumption caused by aircraft performance degradation. In addition, wind tunnel experiment results rarely involve the influence of crosswind on fuel consumption; thus, the results could be inaccurate when compared with field data. This study focuses on the departure climbing phase of aircraft operation and proposes a new fuel consumption model. In this model, the relationships between aerodynamic parameters are extracted by fitting quick access recorder (QAR) actual flight data, and the crosswind effect is also considered. Taking QAR data from two airports in China, the accuracy of the proposed model and its transferability are demonstrated. Applying the proposed model, the fuel saving of a continuous climb operation (CCO) compared with the traditional climb operation is further quantified. Finally, how aircraft mass, climbing angle, and different aircraft models could affect the fuel consumption of the climbing phase of aircraft operation is investigated. The proposed fuel consumption model fills gaps in the existing literature, and the method can be used for developing specific fuel consumption models for more aircraft types at other airports.
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Rahnavard, Mostafa, Moosa Ayati, and Mohammad Reza Hairi Yazdi. "Robust actuator and sensor fault reconstruction of wind turbine using modified sliding mode observer." Transactions of the Institute of Measurement and Control 41, no. 6 (July 19, 2018): 1504–18. http://dx.doi.org/10.1177/0142331218754620.
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This paper proposes a robust fault diagnosis scheme based on modified sliding mode observer, which reconstructs wind turbine hydraulic pitch actuator faults as well as simultaneous sensor faults. The wind turbine under consideration is a 4.8 MW benchmark model developed by Aalborg University and kk-electronic a/s. Rotor rotational speed, generator rotational speed, blade pitch angle and generator torque have different order of magnitudes. Since the dedicated sensors experience faults with quite different values, simultaneous fault reconstruction of these sensors is a challenging task. To address this challenge, some modifications are applied to the classic sliding mode observer to realize simultaneous fault estimation. The modifications are mainly suggested to the discontinuous injection switching term as the nonlinear part of observer. The proposed fault diagnosis scheme does not require know the exact value of nonlinear aerodynamic torque and is robust to disturbance/modelling uncertainties. The aerodynamic torque mapping, represented as a two-dimensional look up table in the benchmark model, is estimated by an analytical expression. The pitch actuator low pressure faults are identified using some fault indicators. By filtering the outputs and defining an augmented state vector, the sensor faults are converted to actuator faults. Several fault scenarios, including the pitch actuator low pressure faults and simultaneous sensor faults, are simulated in the wind turbine benchmark in the presence of measurement noises. Simulation results show that the modified observer immediately and faithfully estimates the actuator faults as well as simultaneous sensor faults with different order of magnitudes.
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Alhmoud, Lina, and Hussein Al-Zoubi. "IoT Applications in Wind Energy Conversion Systems." Open Engineering 9, no. 1 (November 2, 2019): 490–99. http://dx.doi.org/10.1515/eng-2019-0061.
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AbstractRenewable energy reliability has been the main agenda nowadays, where the internet of things (IoT) is a crucial research direction with a lot of opportunities for improvement and challenging work. Data obtained from IoT is converted into actionable information to improve wind turbine performance, driving wind energy cost down and reducing risk. However, the implementation in IoT is a challenging task because the wind turbine system level and component level need real-time control. So, this paper is dedicated to investigating wind resource assessment and lifetime estimation of wind power modules using IoT. To illustrate this issue, a model is built with sub-models of an aerodynamic rotor connected directly to a multi-pole variable speed permanent magnet synchronous generator (PMSG) with variable speed control, pitch angle control and full-scale converter connected to the grid. Besides, a large number of various sensors for measurement of wind parameters are integrated with IoT. Simulations are constructed with Matlab/Simulink and IoT ’Thingspeak’ Mathworks web service. IoT has proved to increase the reliability of measurement strategies, monitoring accuracy, and quality assurance.
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Bianchini, Alessandro, Giovanni Ferrara, Lorenzo Ferrari, and Sandro Magnani. "An Improved Model for the Performance Estimation of an H-Darrieus Wind Turbine in Skewed Flow." Wind Engineering 36, no. 6 (December 2012): 667–86. http://dx.doi.org/10.1260/0309-524x.36.6.667.
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Small turbines are considered one of the most promising technologies for an effective diffusion of renewable energy sources in new installation contexts with a high degree of integration with human activity (e.g. the urban environment). In these new installations, however, the real working conditions can be far from the nominal ones. In particular, the turbine functioning can be noticeably affected by misalignments between the oncoming flow and the axis of the rotor; differently from horizontal-axis wind turbines, whose performance is decreased by a skew angle, H-Darrieus turbines are thought to take advantage from this condition in some cases. In this study, an improved model for the performance prediction of H-Darrieus rotors under skewed flow was developed. In detail, a theoretical approach based on Momentum Models was properly modified to account for the variations induced by the new direction of the flow which invests the rotor. In particular, the modifications in the aerodynamic characteristics of the airfoils, the swept area and the streamtubes distribution were modeled. The performance predictions of the new model were compared both with experimental data available in the technical literature and with the results of wind tunnel tests purposefully carried out on a full scale model of an H-Darrieus turbine. Notable agreement has been constantly obtained between simulations and experiments.
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Копылов, Сергей Николаевич, Леонид Тимофеевич Танклевский, Александр Алексеевич Таранцев, Игорь Александрович Бабиков, and Александр Валерьевич Аракчеев. "Calculated estimation of geometric parameters of automatic water fire extinguishing systems for high rise racks." Pozharnaia bezopasnost`, no. 2(99) (June 18, 2020): 62–69. http://dx.doi.org/10.37657/vniipo.2020.99.2.007.
Повний текст джерелаАнотація:
Рассмотрены вопросы, связанные с применением спринклерных автоматических установок водяного пожаротушения стеллажей. Проанализированы соответствующие нормативные документы. Описаны варианты решения задач определения расхода воды из оросителя, координат места его установки, а также углов распыла огнетушащего вещества и наклона оси оросителя. Приведены примеры расчета геометрических параметров спринклерных автоматических установок водяного пожаротушения стеллажей. Fires at objects whith high-rack storage of combustible materials are particularly dangerous because of the rapid spread of the flame vertically, the risk of collapse of the racks and the damaging effects of high temperature on structural elements of the building. The main method of extinguishing such fires at the initial stage is the use of automatic sprinkler fire extinguishing systems (AUP). The requirements for AUP parameters (types of detectors and sprinklers, their characteristics and distances) depending on the height of the room and storage are currently set out in two normative documents: the set of rules (SP 241.1311500.2015) and the organization standard (VNPB 40-16). Unlike the SP, where there is provided only the supply of a fire extinguishing substance (FES) - water vertically down with high-flow sprinklers of type SOBR (ESFR) and there are no requirements for the type of fire detectors, VNPB provides the use of different types of detectors (aspiration, smoke, heat), forced start-up of AUP sprinklers, which reduces the time of free fire development and the supply of FES by a flow shaper with the spray angle ≈ 600 at an angle to the vertical both to the horizontal and lateral surfaces of the racks. This article discusses the issues of determining the parameters of automatic sprinkler systems for water fire extinguishing of racks. Variants of solving synthesis problems are given - the choice of the places for installing sprinklers depending on the height and width of the racks, their axis of inclination, and also the spray angle. To solve these problems, the computer program called struja.exe was created, a series of calculations on which showed a negligible effect of aerodynamic drag due to relatively small distances. Examples are given. Thus, the features of the sprinkler AUP for the protection of rooms with high-rack storage and the task of determining its geometric parameters are considered. In this case, forced activation of the sprinkler follows in order to avoid a delay in the start of extinguishing. In the future, it is also desirable to conduct additional field experiments with sprinkler water supply and also (if possible) evaluate the effect of ascending flows of combustion products on the water flow from the sprinkler.
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Mitridis, Dimitrios, Chris Bliamis, Pericles Panagiotou, and Kyros Yakinthos. "A novel technique for hypersonic vehicle control." MATEC Web of Conferences 304 (2019): 02008. http://dx.doi.org/10.1051/matecconf/201930402008.
Повний текст джерелаАнотація:
A novel control technique is investigated for hypersonic aerial vehicles. The technique is based on the use of active shock bumps (SBs) as a form of control device. The SBs deflect to create shockwaves on–demand, at specific locations around the aerial vehicle. As a result, a force is applied on the aerial vehicle, which in turn is used to provide the necessary moment for pitch and roll manoeuvres. In this work, a preliminary aerodynamic analysis of the SB device technique is made by means of CFD. For this purpose, and taking the large corresponding Reynolds numbers of the flow into consideration, the two–dimensional Euler equations are solved. A parametric investigation is carried out, by examining the effect of key parameters, namely the Mach number (M) and device deflection angle (δSB) on the produced force acting on the vehicle, serving as a proof of concept. Using a specific interpolation method, the resultant force is presented as a function of the Mach number and the device deflection angle, on three–dimensional charts, where the effect of each parameter is shown (force–Mach–deflection maps). Furthermore, a preliminary feasibility study is performed, including a kinematic analysis and some key material considerations. Additionally, a kinetic analysis is also conducted to secure the dynamic rigidity of the actuating mechanism and provide an initial estimation concerning weight and basic geometrical parameters of the SB mechanism components.
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RUSAK, ZVI, and JANG-CHANG LEE. "Transonic flow of moist air around a thin airfoil with non-equilibrium and homogeneous condensation." Journal of Fluid Mechanics 403 (January 25, 2000): 173–99. http://dx.doi.org/10.1017/s0022112099007053.
Повний текст джерелаАнотація:
A new small-disturbance model for a steady transonic flow of moist air with non-equilibrium and homogeneous condensation around a thin airfoil is presented. The model explores the nonlinear interactions among the near-sonic speed of the flow, the small thickness ratio and angle of attack of the airfoil, and the small amount of water vapour in the air. The condensation rate is calculated according to classical nucleation and droplet growth models. The asymptotic analysis gives the similarity parameters that govern the flow problem. Also, the flow field can be described by a non-homogeneous (extended) transonic small-disturbance (TSD) equation coupled with a set of four ordinary differential equations for the calculation of the condensate (or sublimate) mass fraction. An iterative numerical scheme which combines Murman & Cole's (1971) method for the solution of the TSD equation with Simpson's integration rule for the estimation of the condensate mass production is developed. The results show good agreement with available numerical simulations using the inviscid fluid flow equations. The model is used to study the effects of humidity and of energy supply from condensation on the aerodynamic performance of airfoils.
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Gologan, C., K. Broichhausen, and J. Seifert. "A calculation method for parametric design studies of V/STOL aircraft." Aeronautical Journal 113, no. 1143 (May 2009): 309–17. http://dx.doi.org/10.1017/s0001924000002980.
Повний текст джерелаАнотація:
Abstract This paper provides a method that helps the aircraft designer to develop a performance constraint chart (PCC) for vertical/short takeoff and landing (V/STOL) aircraft that produce hybrid lift (static lift in combination with aerodynamic lift). The PPC provides a first estimation for the thrust-to-weight ratio (T 0/MTOW) and wing loading (MTOW/S). The method is applicable to concepts, where static lift and main-engine thrust are coupled (e.g. the F-35B system, turbojet and lift-fan, coupled by a shaft) and to concepts, where static lift is produced by separate devices (e.g. lift-engines or other concepts for static lift). It includes thrust vectoring of the main engine. The method includes the evaluation of certain flight stages, or segments, as short take-off and landing (STOL), vertical landing (VL), one engine inoperative (OEI) climb (for civil aircraft concept applications) and cruise, all these have to be considered. For each of these five segments, standard flight mechanics equations are extended by a static lift component, an augmentation ratio (a factor that describes the dependency of the thrust and the static lift, if coupled) and the thrust vectoring angle. Hence these equations are modified in a way that the aircraft designer can directly calculate T0/MTOW and MTOW/S for the performance requirements of each segment. Thus, an optimum design point can be selected. Inputs are aerodynamic coefficients, maximum lift coefficient of the wing, mass fraction from take-off to landing, additional static lift during take-off and landing, number of engines, augmentation ratio of the propulsion system and required take-off and landing field length. A performance constraint chart for a JSF F-35B type aircraft is modelled to show the application of the method. As an application to civil aircraft a PCC and a parameter optimisation for the civil regional jet ‘HyLiner’ is presented.