Validation of computed high-lift flows with significant wind-tunnel effects

Nonlinear longitudinal aerodynamics associated with cascade fins at high angles of attack near stall has been modelled using Kirchhoff's formulation. Grid fins are a relatively recent development in guided missile technology. In this paper, a new category of grid fins, nomenclatured as cascade fins, has been proposed. In cascade fin design, an appropriate selection of gap-to-chord ratio and the number of planar members lead to desired stall angle and acceptable overall lift coefficient, respectively. Kirchhoff's steady-state stall model has been validated on wind tunnel data generated for Cascade fins having rectangular airfoil cross-section. National Wind Tunnel Facility (NWTF) of IIT, Kanpur, was used to generate the wind tunnel data consisting of the variation of lift coefficient with angle of attack. The cascade fins were tested to generate the data by varying gap-to-chord ratio and number of planar fins. The cascade fins with rectangular cross-section were tested with and without end plates. Kirchhoff's steady-state stall model was applied to wind tunnel data of cascade fins for modelling flow separation point and maximum likelihood method was used to estimate the parameters characterising stall characteristics. The effects of end plates, variation of number of fins and gap-to-chord ratio on parameter estimation were also studied. It has been observed that Kirchhoff's steady-state stall model could advantageously be applied to model nonlinear aerodynamics associated with cascade fins at high angle of attack. Defence Science Journal, 2011, 61(2), pp.157-164 , DOI:http://dx.doi.org/10.14429/dsj.61.481

At present, there is a great interest in the development of new airfoils for wind turbines and high-lift wings of unmanned aerial vehicles (UAV). The requirements for such airfoils differ from conventional aircraft airfoils, because of structural reasons and extreme operating conditions. So, wind turbine airfoils operate frequently under fully separated flow when stall is used for power regulation at high wind speeds. At the same time design of airfoils for wings UAV poses the problem of availability of high-lift at low Reynolds number. Modern airfoils are to a large extent developed from numerical methods. However, the complex flow conditions such as separation at high angles of attack, laminar separation bubbles and the transition from laminar to turbulent flow are difficult to predict accurately. Hence, testing of airfoils at a two-dimensional condition is an important phase in airfoil design. The development and validation of a 2D testing facility for investigation of single and multi-element airfoils in the wind tunnel Т-102 with open test section are considered in this article. T-102 is a continuous-operation, closed-layout wind tunnel with two reverse channels. The test section has an elliptical cross-section of 4 ×2,33 m and a length of 4 m. Two big flat panels of the L × H=3 ×3,9 m size installed upright on balance frame aligned with the free stream are used for simulating two-dimensional flow in the tunnel test section. The airfoil section in the layout of a rectangular wing is mounted horizontally between flat panels with minimum gaps to ensure 2D flow conditions. The aerodynamic forces and pitch moment acting on the model were measured by wind tunnel balance. To determine boundary corrections for a new test section of wind tunnel, the experimental investigation of three geometrically similar models has been executed. The use of boundary corrections has provided good correlation of the test data of airfoil NACA 6712 with the results obtained from the wind tunnel except for lift and drag coefficient values at high angles of attack.

In this paper, an intelligent algorithm for the aerodynamic parameter calibration of wind tunnel experiments of advanced layout aircraft at a high angle of attack is proposed. The proposed algorithm is based on a homologous comparison and tuning algorithm, and it can effectively improve the accuracy of the wind tunnel experiment model. First, based on the analysis of large oscillation wind tunnel experimental data of an advanced layout scaled aircraft, the high angle of attack wind tunnel experimental model composed of static derivative, dynamic derivative, and rotating balance derivative is established. Second, to improve the accuracy of the wind tunnel experimental model effectively, the idea of combing layered calibration and intelligent algorithm for high angle of attack homologous comparison correction is proposed. The proposed method solves the problems of complex structure, a large amount of data, and poor accuracy in homologous comparison of high angle of attack aerodynamic models of advanced layout aircraft. Finally, the homologous comparison interface software is designed based on MATLAB GUI, which integrates the proposed methods and ideas and realizes effective adjustment of aerodynamic parameters of high angle of attack simulation flight wind tunnel test of an advanced layout aircraft. This study provides a reliable engineering and technical means for subsequent high angle of attack flight test verification of the advanced layout aircraft.

This paper presents the development of a nonlinear aerodynamic model that describes the impact of forebody tangential blowing (FTB) on the generated roll moment for a fixed aircraft attitude. This model uses Nonlinear Indicial Responses (NlR) and incorporates physical insight gained from experimental data. The model accounts for the nonlinear aerodynamic behavior observed when FTB is applied at high angles of attack. Its predictions for the roll moment generated by FTB are compared to actual data and a model that uses a tirstorder lag of the static loads. The NlR model is found to provide a better prediction of the measured transient response. Experimental results used in the model’s development are discussed. They include flow visualization as well as static and dynamic measurements of the aerodynamic loads for various roll angles and blowing mass flow rates. The experimental apparatus used in this study is described and consists of an FTB-equipped wind tunnel model which is free to roll. The model is a wing-body combination with slots on the forebody for tangential blowing and is nominally set at 45 degrees angle of attack. To augment the control of a flight vehicle at high angles of attack, several methods of active flow control have been proposed and explored.4*5 They provide a means to alter the flow structure in a rational manner and increase the controllability of the vehicle under these flight conditions. The injection of thin sheets of air tangentially to the forebody of the vehicle has been found to be an extremely promising method6-8 and progress has been made in demonstrating the feasibility of using Forebody Tangential Blowing (FTB) to control the roll-yaw motion of flight vehicles through both experimental demonstrationsg*10 and simulations.” Experimental investigations into the use of FTB have been conducted at the low-speed wind tunnel of the Department of Aeronautics and Astronautics at Stanford University. The experimental apparatus consists of a wind tunnel model that is free to ro11.‘2 The model has a cone-cylinder fuselage and a sharp leading-edge delta wing. It is set at a nominal incidence angle of 45 degrees and is equipped with side slots to support the injection of a thin sheet of air tangentially to the forebody of the vehicle.

As part of a project to determine ship air wake impact on naval rotary wing vehicles, this paper compares wind tunnel and computational fluid dynamics (CFD) simulations with in situ data that has been collected on an instrumented United States Naval Academy (USNA) YP (Patrol Craft, Training). The in situ data has been collected underway on a vessel which is 108 ft long and has a 24 ft above water line height. The wind tunnel data is from a 4% scale model tested in the USNA 42 × 60 × 120 inch wind tunnel. CFD simulations were performed on a 15.5 million tetrahedral unstructured grid. Data from the three sources is compared over 11 vertical and horizontal planes above the flight deck on the stern section of the YP. Data is compared for a head wind condition and for relative wind 15° off the starboard bow. Good comparison in velocity direction, typically less than 10-15°, is observed in the vertical and horizontal planes. Normalized velocity magnitudes do not compare as well, likely arising from the lack of an atmospheric boundary layer modeled in CFD simulations and wind tunnel testing.

Upper surface flaps commonly referred to as spoilers or drag brakes can increase maximum lift, and improve aerodynamic efficiency at high, near-stall angles of attack. This phenomenon was studied experimentally and computationally using a 0.307626 m chord length NACA 2412 airfoil in six different configurations, and one baseline clean configuration. A wind tunnel model was placed in the Ryerson Low Speed Wind Tunnel (atmospheric, closed-circuit, 3 ft × 3 ft test section) at a Reynold’s number of approximately 780,000 and a Mach number of 0.136. The wind tunnel study increased the lift coefficient by 0.393%-2.497% depending on the spoiler configuration. A spoiler of 10% chord length increased the maximum lift coefficient by 2.497 % when deflected 8º, by 2.110% when deflected 15º, and reduced the maximum lift coefficient by 2.783% when deflected 25º. A spoiler of 15% chord length produced smaller maximum lift coefficient gains; 0.393% when deflected 8º, by 1.760% when deflected 15º, and reduced the maximum lift coefficient by 4.475% when deflected 25º. Deflecting the spoiler increased the stall angle between 37.658% and 87.544% when compared with the clean configuration. The drag coefficient of spoiler configurations was lower than the clean configuration at angles of attack above 18º. The combination of the increased lift and reduced drag at angles of attack above 18º created by the spoiler configurations resulted in a higher aerodynamic efficiency than the clean configuration case. A 10% chord length spoiler deflected at 8º produced the highest aerodynamic efficiency gains. At low angles of attack, the computational study produced consistently higher lift coefficients compared with the wind tunnel experiment. The lift-slope was consistent with the wind tunnel experiment lift-slope. The spoiler airfoil stall behaviour was inconsistent with the results from the wind tunnel experiment. The drag coefficient results were consistent with the wind tunnel experiment at low angles of attack. However, the spoiler equipped airfoils did not reduce drag at high angles of attack. Therefore, the computational model was not valid for the spoiler configurations at high angles of attack.

Computational and experimental results are presented for delta wings with vortex flaps. The wings have an undeflected leading-edge sweep of 75 deg. Flap angles of 5 deg and 10 deg, measured in the streamwise direction, are considered. The nominal angle of attack alpha is varied from 4 deg to 12 deg. Results for freestream Mach numbers of 1.7 and 2.4 are shown. Surface pressure, tuft patterns and vapor screens are given for the experimental data. Surface pressure, tuft patterns, cross-flow velocities, total pressure loss and cross-flow Mach number are given for the numerical data. C(l)-vs-alpha curves are shown for experimental and computational results. The flows are shown to be very sensitive to angle of attack, and the agreement between experimental and theoretical results is improved if the calculations are made at angles of attack slightly larger than the experimental angles of attack. The computational model correctly predicts the topology of the flow in each of the cases considered. The lift is predicted well at the higher angles of attack, but slightly overpredicted at the lower angles of attack.

The present study investigates the differences in the predicted lift for a wind tunnel model from either surface pressures or side wall pressures. The investigated model is the NACA63-418 tested in the Poul la Cour tunnel. The measurements are compared with previous conducted wind tunnel campaigns for a Reynolds number of 3E6. The 3D CFD model of the wind tunnel will be used for investigating details of the flow mechanisms, and to validate assumptions and findings from the experimental data. The CFD simulations show that when the flow is attached the lift based on the side wall pressures results in almost the same lift as calculated from the surface pressures. Around the maximum lift and in the stalled region the flow is not anymore 2D. For the NACA63-418 airfoil the lift based on the side wall pressures are generally highest. The CFD simulations confirm that the method used to predict the lift from the limited side wall pressure strips in the wind tunnel is applicable.

A procedure for the estimation of airplane model structure and parameters is applied to data from a modern fighter airplane operating within an angle of attack range of 5 to 60 deg. The paper briefly describes the airplane, flight and wind tunnel data available, and the estimation method. The results presented contain basic longitudinal characteristics of the airplane and the estimates of aerodynamic parameters in the yawing-moment equations. These estimates are obtained from small and large amplitude maneuvers. Because the latter set of data was not suitable for airplane identification, some of the large amplitude maneuvers were joined together and then partitioned into subsets according to the values of angle of attack. Each subset was then analyzed as a separate data set. Most of the estimated parameters and functions are in good agreement with the wind tunnel measurements. The estimated lateral parameters in the model equations also demonstrate good prediction capabilities.

Flow visualization experiments were performed on a series of generic fighter configurations at medium to high angles of attack in both a w ater tunnel and a w ind tunnel. The models were tested originally in a w ater tunnel to investigate methods for controlling the forebody and wing/LEX vortices for airframe control augmentation at high angles of attack. The same models were subsequently tested in a low-speed wind tunnel at the identical Reynolds number to compare the core locations and burst points of the LEX and wing vortices in the two test mediums. A u nique application of a smoke flow visualization method used primarily in low-speed wind tunnels for two-dimensional airfoil flow visualization was adapted to provide a direct means for comparison to the dye-flow results from the water tunnel. Direct comparisons of photographs from the two tests at various angles of attack on three different configurations showed nearly identical flow fields, confirming expectations that the test medium should have n o bearing on the flow field structure.

Slender-body aircraft operating at high angles of attack often experience nonlinear, asymmetric multi-vortex flow structures that generate random, unsteady lateral forces, undermining stability, and maneuverability. Dielectric barrier discharge plasma actuators can eliminate these lateral forces. However, conventional open-loop plasma control method cannot adapt to dynamic flow fields in real time, limiting the overall effectiveness of active flow control. This study introduces a plasma control framework grounded in physical principles and develops plasma actuator design methods to regulate vortex interactions, stabilize flow dynamics, and optimize control efficiency. An intelligent closed-loop flow control strategy based on proximal policy optimization, a deep reinforcement learning algorithm, is utilized to enable real-time plasma parameter adjustments for suppressing lateral force at high angle of attack. The spatiotemporal interaction of plasma-induced and asymmetric vortices was investigated through synchronized pressure measurements and particle image velocimetry. The proximal policy optimization-based parameter optimization model was trained online in an educational open-return wind tunnel and subsequently deployed in a low-speed closed-return wind tunnel. Based on vortex stability analysis and comprehensive results, the closed-loop control algorithm significantly mitigates lateral forces, achieving a 68.5% reduction compared to steady plasma actuation, while improving energy efficiency by 70% over conventional methods.

<div class="htmlview paragraph">Wind tunnel interference effects are still considered to be negligible - or at least undesired - in automotive aerodynamics. Consequently, up to now there is no standard correction method which is used in everyday wind tunnel testing although a lot of research has been done in recent years.</div> <div class="htmlview paragraph">In most full-vehicle CFD simulations, wind tunnel interference effects are not simulated. The flow about the car is computed under idealized conditions. The wind tunnel is designed to simulate these conditions but fails to do so to some degree due to its limited size.</div> <div class="htmlview paragraph">Therefore a comparison of blockage-free CFD results and wind tunnel measurements is deficient. Hence CFD simulations including wind tunnel interference effects should be favored in the future for validation purposes. Furthermore, CFD offers new possibilities to investigate individual contributions to wind tunnel interference effects and therefore could help to increase the understanding of the flow in the wind tunnel. This may result in the development of advanced correction methods. To be able to use these possibilities, it has to be ensured that all effects of the whole wind tunnel test section with a car present will be captured in a simulation.</div> <div class="htmlview paragraph">In this paper simulations of the IVK model scale wind tunnel including wind tunnel interference effects are presented. As a first parameter variation, the influence of different longitudinal pressure distributions in the virtual test section are investigated and compared with test data. This investigation is presented here. In a second step, different vehicle models will be tested. All simulations were carried out with EXA PowerFLOW™.</div>

A video image analysis system for concentration measurements and flow visualization in building wakes

Wind tunnel simulation of airflow and pollutant diffusion over complex terrain

Comparison of aerodynamic roughness measured in a field experiment and in a wind tunnel simulation