Simulation of a passenger aircraft flight with the wing tip cut

2021 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Zdobyslaw Jan Goraj ◽  
Mariusz Kowalski ◽  
Łukasz Kiszkowiak ◽  
Aleksander Olejnik
Keyword(s):  
Wind Tunnel ◽  
Flow Separation ◽  
Pitch Angle ◽  
Equations Of Motion ◽  
Aerodynamic Characteristics ◽  
Final Phase ◽  
Total Flow ◽  
Content Type ◽  
Passenger Aircraft ◽  
Wing Tip

Purpose The purpose of this paper is to present the result of simulations that were performed to assess the uncontrolled motion of the passenger aircraft following its wing tip was suddenly cut. Such a simulation can help to understand the tendencies of aircraft behaviour after wing tip cut, assess how fast this phenomenon is going on and estimate the values of angles of attack, sideslip and pitch angle basing on given aerodynamic characteristics. Also, answer the question if pilot can counteract high deviations from flight path initially planned during the final phase of approach to landing. Design/methodology/approach Simulation is based on the full non-linear equations of motion derived from generalised equations of change of momentum and moment of momentum of rigid body. Dynamic equations of motion in the so-called normal mode are solved in the so-called stability frame of reference. Findings It was found that asymmetric rolling moment must be compensated by essential increase of pitching moment. Moreover, it appeared that aircraft goes into high angles of attack and high pitch angle and, therefore, for reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs. Practical implications For accurate simulation, it is strongly recommended to perform to perform first the wind tunnel testing in the range of +20o ÷ 120o and use it in flight simulation. Originality/value The presented methodology is an original for numerical simulation of flight trajectory during the final phase of approach to landing in a hazardous state of flight. For reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs, whereas usually maximum angles of attack used in wind tunnel experiments for passenger aircraft are not higher than 25 degrees. The influence of limited range of experimental data on results of simulation is another value which can be adopted in the future investigations of hazardous states of flight.

An experimental parametric study on planar sheared wings tip devices for low-to-moderate aspect ratio wings

2021 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Lourelay Moreira dos Santos ◽  
Guilherme Ferreira Gomes ◽  
Rogerio F. Coimbra
Keyword(s):  
Wind Tunnel ◽  
Aspect Ratio ◽  
Parametric Study ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Sweep Angle ◽  
Wind Tunnel Tests ◽  
Content Type ◽  
Wing Tips ◽  
Wing Tip

Purpose The purpose of this study is to investigate the aerodynamic characteristics of a low-to-moderate-aspect-ratio, tapered, untwisted, unswept wing, equipped of sheared wing tips. Design/methodology/approach In this work, wind tunnel tests were made to study the influence in aerodynamic characteristics over a typical low-to-moderate-aspect-ratio wing of a general aviation aircraft, equipped with sheared – swept and tapered planar – wing tips. An experimental parametric study of different wing tips was tested. Variations in its leading and trailing edge sweep angle as well as variations in wing tip taper ratio were considered. Sheared wing tips modify the flow pattern in the outboard region of the wing producing a vortex flow at the wing tip leading edge, enhancing lift at high angles of attack. Findings The induced drag is responsible for nearly 50% of aircraft total drag and can be reduced through modifications to the wing tip. Some wing tip models present complex geometries and many of them present benefits in particular flight conditions. Results have demonstrated that sweeping the wing tip leading edge between 60 and 65 degrees offers an increment in wing aerodynamic efficiency, especially at high lift conditions. However, results have demonstrated that moderate wing tip taper ratio (0.50) has better aerodynamic benefits than highly tapered wing tips (from 0.25 to 0.15), even with little less wing tip leading edge sweep angle (from 57 to 62 degrees). The moderate wing tip taper ratio (0.50) offers more wing area and wing span than the wings with highly tapered wing tips, for the same aspect ratio wing. Originality/value Although many studies have been reported on the aerodynamics of wing tips, most of them presented complex non-planar geometries and were developed for cruise flight in high subsonic regime (low lift coefficient). In this work, an exploration and parametric study through wind tunnel tests were made, to evaluate the influence in aerodynamic characteristics of a low-to-moderate-aspect-ratio, tapered, untwisted, unswept wing, equipped of sheared wing tips (wing tips highly swept and tapered).

Fluidic actuators for separation control at the engine/wing junction

2017 ◽  
Vol 89 (5) ◽  
pp. 709-718 ◽  
Author(s):  
Philipp Schloesser ◽  
Michael Meyer ◽  
Martin Schueller ◽  
Perez Weigel ◽  
Matthias Bauer
Keyword(s):  
Wind Tunnel ◽  
Flow Control ◽  
Flow Separation ◽  
Mass Flux ◽  
Large Scale ◽  
Active Flow Control ◽  
Wind Tunnel Test ◽  
Local Flow ◽  
Content Type ◽  
Active Flow

Purpose The area behind the engine/wing junction of conventional civil aircraft configurations with underwing-mounted turbofans is susceptible to local flow separation at high angles of attack, which potentially impacts maximum lift performance of the aircraft. This paper aims to present the design, testing and optimization of two distinct systems of fluidic actuation dedicated to reduce separation at the engine/wing junction. Design/methodology/approach Active flow control applied at the unprotected leading edge inboard of the engine pylon has shown considerable potential to alleviate or even eliminate local flow separation, and consequently regain maximum lift performance. Two actuator systems, pulsed jet actuators with and without net mass flux, are tested and optimized with respect to an upcoming large-scale wind tunnel test to assess the effect of active flow control on the flow behavior. The requirements and parameters of the flow control hardware are set by numerical simulations of project partners. Findings The results of ground test show that full modulation of the jets of the non-zero mass flux actuator is achieved. In addition, it could be shown that the required parameters can be satisfied at design mass flow, and that pressure levels are within bounds. Furthermore, a new generation of zero-net mass flux actuators with improved performance is presented and described. This flow control system includes the actuator devices, their integration, as well as the drive and control electronics system that is used to drive groups of actuators. Originality/value The originality is given by the application of the two flow control systems in a scheduled large-scale wind tunnel test.

Wind tunnel tests of the inverted joined wing and a comparison with CFD results

2018 ◽  
Vol 90 (4) ◽  
pp. 586-601
Author(s):  
Cezary Galinski ◽  
Grzegorz Krysztofiak ◽  
Marek Miller ◽  
Pawel Ruchala ◽  
Marek Kalski ◽  
...  
Keyword(s):  
Wind Tunnel ◽  
Aerodynamic Characteristics ◽  
Control Surface ◽  
Dual Use ◽  
Wind Tunnel Tests ◽  
Content Type ◽  
Flying Qualities ◽  
Computational Fluid Dynamics Cfd ◽  
Joined Wing ◽  
Surface Deflections

Purpose The purpose of this paper is to present the methodology and approach adapted to conduct a wind tunnel experiment on the inverted joined-wing airplane flying model together with the results obtained. Design/methodology/approach General assumptions underlying the dual-use model design are presented in this paper. The model was supposed to be used for both wind tunnel tests and flight tests that significantly drive its size and internal structure. Wind tunnel tests results compared with the outcome of computational fluid dynamics (CFD) were used to assess airplane flying qualities before the maiden flight was performed. Findings Extensive data about the aerodynamic characteristics of the airplane were collected. Clean configurations in symmetric and asymmetric cases and also configurations with various control surface deflections were tested. Practical implications The data obtained experimentally made it possible to predict the performance and stability properties of the unconventional airplane and to draw conclusions on improvements in further designs of this configuration. Originality/value The airplane described in this paper differs from frequently analyzed joined-wing configurations, as it boasts a front lifting surface attached at the top of the fuselage, whereas the aft one is attached at the bottom. The testing technique involving the application of a dual-use model is also innovative.

Flow Separation and Stall Characteristics of Plane, Constant-Section Wings in Subcritical Flow

1968 ◽  
Vol 72 (685) ◽  
pp. 82-90 ◽  
Author(s):  
P. D. Chappell
Keyword(s):  
Flow Separation ◽  
Lift Coefficient ◽  
Aerodynamic Characteristics ◽  
Empirical Correlations ◽  
Trailing Edge ◽  
Subcritical Flow ◽  
Mach Numbers ◽  
Wing Tip ◽  
Constant Section

Summary This paper discusses the various aspects of flow separation on wings and aerofoils at subcritical Mach numbers. It is shown that in general the onset of non-linearities in the aerodynamic characteristics, for plane constant-section wings, can be related to the initiation of regions of flow separation somewhere on the wing and often near the wing tip. Empirical correlations are produced, tor plane wings of constant symmetrical section, enabling estimates to be made of the lift coefficient for onset of separation at either the leading or trailing edge of the wing. A correlation of stall types on aerofoils, due to Gault, is extended in application to the estimation of initial stall and separation types for plane wings of constant symmetrical section. Suggestions for extending the application of this correlation to the estimation of section stall type for cambered, twisted and non-constant section wings are made.

An integral approach for aircraft pitch control and instrumentation in a wind-tunnel

2020 ◽  
Vol 92 (7) ◽  
pp. 1111-1123
Author(s):  
José Francisco Villarreal Valderrama ◽  
Luis Takano ◽  
Eduardo Liceaga-Castro ◽  
Diana Hernandez-Alcantara ◽  
Patricia Del Carmen Zambrano-Robledo ◽  
...  
Keyword(s):  
Wind Speed ◽  
Wind Tunnel ◽  
Pitch Angle ◽  
Test Bench ◽  
Operating Conditions ◽  
Experimental Methodology ◽  
Control Algorithms ◽  
Content Type ◽  
Pitch Control ◽  
Model Free

Purpose Aircraft pitch control is fundamental for the performance of micro aerial vehicles (MAVs). The purpose of this paper is to establish a simple experimental procedure to calibrate pitch instrumentation and classical control algorithms. This includes developing an efficient pitch angle observer with optimal estimation and evaluating controllers under uncertainty and external disturbances. Design/methodology/approach A wind tunnel test bench is designed to simulate fixed-wing aircraft dynamics. Key elements of the instrumentation commonly found in MAVs are characterized in a gyroscopic test bench. A data fusion algorithm is calibrated to match the gyroscopic test bench measurements and is then integrated into the autopilot platform. The elevator-angle to pitch-angle dynamic model is obtained experimentally. Two different control algorithms, based on model-free and model-based approaches, are designed. These controllers are analyzed in terms of parametric uncertainties due to wind speed variations and external perturbation because of sudden weight distribution changes. A series of experimental tests is performed in wind-tunnel facilities to highlight the main features of each control approach. Findings With regard to the instrumentation algorithms, a simple experimental methodology for the design of optimal pitch angle observer is presented and validated experimentally. In the context of the platform design and identification, the similitude among the theoretical and experimental responses shows that the platform is suitable for typical pitch control assessment. The wind tunnel experiments show that a fixed linear controller, designed using classical frequency domain concepts, is able to provide adequate responses in scenarios that approximate the operation of MAVs. Research limitations/implications The aircraft orientation observer can be used for both pitch and roll angles. However, for simultaneousyaw angle estimation the proposed design method requires further research. The model analysis considers a wind speed range of 6-18 m/s, with a nominal operation of 12 m/s. The maximum experimentally tested reference for the pitch angle controller was 20°. Further operating conditions may require more complex control approaches (e.g. scheduling, non-linear, etc.). However, this operating range is enough for typical MAV missions. Originality/value The study shows the design of an effective pitch angle observer, based on a simple experimental approach, which achieved locally optimum estimates at the test conditions. Additionally, the instrumentation and design of a test bench for typical pitch control assessment in wind tunnel facilities is presented. Finally, the study presents the development of a simple controller that provides adequate responses in scenarios that approximate the operation of MAVs, including perturbations that resemble package delivery and parametric uncertainty due to wind speed variations.

Research and selection of MALE wing profile

2019 ◽  
Vol 91 (2) ◽  
pp. 264-271 ◽  
Author(s):  
Cezary Gorniak ◽  
Zdobyslaw Jan Goraj ◽  
Bartosz Olszanski
Keyword(s):  
Wind Tunnel ◽  
Early Stage ◽  
Phase 1 ◽  
Lift Coefficient ◽  
High Volume ◽  
Aerodynamic Characteristics ◽  
Preliminary Design ◽  
Content Type ◽  
Wing Section ◽  
Selection Of

Purpose The purpose of this research is a preliminary selection of wing section, which would be the best suited for PW-100 – a MALE class UAV of 600 kg weight. PW-100 will be used as a testing platform in different institutions such as research institutes, industry research centers or universities of technology (phase 1) to enable the in-flight testing of various on-board systems (mobile radars, thermovision sensors, chemical sensors, antennas, teledetection systems and others). Untypical layout of PW-100 resulted from the plans of further development of this configuration for a military application. Design/methodology/approach Important role in the research described in this paper plays the selection of main wing section to fulfil the preliminary requirements regarding maximum lift coefficient, minimum drag, aerodynamic efficiency etc. Two different wing sections (R1082 and SA19) were tested in wind tunnel, both with flaps deflected at the range of 0°-30°. Experimental measurements were performed in the low turbulence wind tunnel with closed test section of 45 cm × 35 cm. Numerical simulations of the flow around the wing sections were performed using MSES code. Boundary conditions were assumed basing on the typical mission of PW-100 for flight altitude around 9,000 m, speed of 110 km/h what results in Re = 956,000. Findings Lift coefficients obtained from both experimental and numerical methods for single slatted airfoil SA19 are much higher than that of get for Ronch R1082 airfoil. PW-100 aircraft with SA19 airfoils will be able to be trimmed and fly at any altitude up to 9,000 m and with an arbitrary weight up to 600 kg. Aerodynamic characteristics of SA19 remain smoother and more predictable than that of R1082 airfoil. The very promising properties of SA19 airfoil are well known to the authors since the beginning of last decade when PW team worked together with IAI team on CAPECON project and now it was fully confirmed by this research. Practical implications It was confirmed that selection of the proper wing section for the special mission performed by UAV is of the highest importance decision to be taken at the preliminary design phase. Because there is a limited access to the base of technical parameters in many different UAVs classes and the classical analysis of trends cannot be fully applied, the wing section analysis, either experimental or numerical, must be performed. The situation is much worse than in the case of manned aircrafts because most of the modern UAVs are made of carbon or glass fiber, and therefore, there is no chance for analysis of trends. Originality/value This paper presents a very efficient method of assessing the influence of wing section on aircraft performance adopted for MALE class UAV, especially in an early stage of preliminary design process. The assessment is built mainly on three requirements: Maximum 2D lift coefficient for take-off configuration with flap deflected on 20 degrees should be greater than 2.4. Endurance factor CL1.5/CD for loitering conditions (Ma = 0.5 and CD0 = 0.008) should be greater than 110. The relative wing section thickness should be greater/equal than 19 per cent (it is required for high volume fuel tank located in the wings).

Supersonic Wind Tunnel Experiment on Aerodynamic Characteristics and Winglets Effects of the Tapered Supersonic Biplane

2011 ◽  
Author(s):  
Takashi Fujisono ◽  
Hiroshi Yamashita ◽  
Atsushi Toyoda ◽  
Hiroki Nagai ◽  
Keisuke Asai ◽  
...  
Keyword(s):  
Wind Tunnel ◽  
High Speed ◽  
Type A ◽  
Aerodynamic Characteristics ◽  
Type B ◽  
Supersonic Wind Tunnel ◽  
Wing Model ◽  
Mach Numbers ◽  
The Difference ◽  
Wing Tip

The aerodynamic characteristics and the effects of tip plates of a tapered supersonic biplane wing during the starting process have been investigated through Experimental and Computational Fluid Dynamics (EFD/CFD). Three types of the wing model were used: without tip plate (type-N); with the tip plate which covers only the aft-half of the wing tip (type-A); with the tip plate which covers the entire wing tip (type-B). Experiment was conducted in the supersonic blowdown wind tunnel with 600 mm × 600 mm cross section located at the High-speed Wind Tunnel Facility of Institute of Space and Astronautical Science (ISAS/JAXA). The flow conditions covered from M∞ = 1.5 to 1.9 with increments of 0.1. Pressure-Sensitive Paint was applied to measure pressure distributions on the surface of the wing. CFD simulations were conducted to compare with experiments and to investigate effects of the Mach numbers in detail. The tapered biplane wing without the tip plate was found to start between M∞ = 1.8 and 1.9. The difference of the starting Mach numbers between type-N and type-A was small. On the other hand, the starting Mach number of type-B was about 0.05 higher than that of type-N.

scholarly journals Konsistensi dan Akurasi Data Hasil Pengukuran pada Pengujian Aerodinamika Model Pesawat Jenis Penumpang Sipil (AirLiner) di Wind Tunnel BBTA3-BPPT

2019 ◽  
Vol 2 (1) ◽  
Author(s):  
Novan Risnawan ◽  
F. Andree Yohanes ◽  
Sunarno Sunarno ◽  
Alief Sadlie Kasman
Keyword(s):  
Wind Tunnel ◽  
Design Process ◽  
National Laboratory ◽  
Aircraft Design ◽  
Aerodynamic Characteristics ◽  
Data Consistency ◽  
Aircraft Model ◽  
Measurement Results ◽  
Passenger Aircraft ◽  
Test Result

AbstractAerodynamic testing of an aircraft model in wind tunnel is important step in an aircraft design process. National Laboratory of Aerodynamic, Aeroelastic and Aeroacoustics Technology (BBTA3) have conducted a testing of a type of civilian passenger aircraft model to obtain aerodynamic characteristics data in ILST wind tunnel. The test result will be used in design process to predict performance and stability, so that the data must be consistent and accurate. This paper discusses the process of data consistency and accuracy checking by performing short repeatability and long repeatability testing so that the measurement results can be regarded valid and can be used to represent the validity of whole measurement data.Keywords: Consistency of test result data, wind tunnel, aerodynamics.Abstraks Pengujian aerodinamik sebuah model pesawat di terowongan angin merupakan tahap penting dalam proses desain pesawat. Balai Besar Teknologi Aerodinamika, Aeroelastika dan Aeroakustika (BBTA3) telah melakukan pengujian sebuah model pesawat bertipe penumpang sipil (airliner) untuk mendapatkan data karakteristik aerodinamika di terowongan angin ILST. Data hasil pengujian ini akan digunakan di dalam proses disain untuk memprediksi kinerja dan kestabilan pesawat sehingga data tersebut harus konsisten dan akurat. Tulisan ini mendiskusikan proses pengecekan konsistensi dan akurasi data dengan melakukan pengujian keberulangan dekat dan pengujian keberulangan jauh sehingga hasil pengukuran dapat dinyatakan valid dan dapat digunakan untuk mewakili keabsahan seluruh data pengukuran.Kata Kunci : Konsistensi data hasil pengujian, wind tunnel, aerodinamika.

Numerical studies of active flow control on wing tip extension

2019 ◽  
Vol 91 (2) ◽  
pp. 346-352
Author(s):  
Petr Vrchota ◽  
Ales Prachar ◽  
Shia-Hui Peng ◽  
Magnus Tormalm ◽  
Peter Eliasson
Keyword(s):  
Flow Control ◽  
Mass Flow ◽  
Flow Separation ◽  
Physical Modeling ◽  
Active Flow Control ◽  
Leading Edge ◽  
Content Type ◽  
Tip Extension ◽  
Active Flow ◽  
Wing Tip

Purpose In the European project AFLoNext, active flow control (AFC) measures were adopted in the wing tip extension leading edge to suppress flow separation. It is expected that the designed wing tip extension may improve aerodynamic efficiency by about 2 per cent in terms of fuel consumption and emissions. As the leading edge of the wing tip is not protected with high-lift device, flow separation occurs earlier than over the inboard wing in the take-off/landing configuration. The aim of this study is the adoption of AFC to delay wing tip stall and to improve lift-to-drag ratio. Design/methodology/approach Several actuator locations and AFC strategies were tested with computational fluid dynamics. The first approach was “standard” one with physical modeling of the actuators, and the second one was focused on the volume forcing method. The actuators location and the forcing plane close to separation line of the reference configuration were chose to enhance the flow with steady and pulsed jet blowing. Dependence of the lift-to-drag benefit with respect to injected mass flow is investigated. Findings The mechanism of flow separation onset is identified as the interaction of slat-end and wing tip vortices. These vortices moving toward each other with increasing angle of attack (AoA) interact and cause the flow separation. AFC is applied to control the slat-end vortex and the inboard movement of the wing tip vortex to suppress their interaction. The separation onset has been postponed by about 2° of AoA; the value of ift-to-drag (L/D) was improved up to 22 per cent for the most beneficial cases. Practical implications The AFC using the steady or pulsed blowing (PB) was proved to be an effective tool for delaying the flow separation. Although better values of L/D have been reached using steady blowing, it is also shown that PB case with a duty cycle of 0.5 needs only one half of the mass flow. Originality/value Two approaches of different levels of complexity are studied and compared. The first is based on physical modeling of actuator cavities, while the second relies on volume forcing method which does not require detailed actuator modeling. Both approaches give consistent results.

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