Assessment of a Data Assimilation Technique for Wind Tunnel Wall Interference Corrections

Aerodynamic characteristics of a guided rocket projectile are numerically simulated. Simulations are performed in a range of flight conditions using a commercial CFD software FLUENT, and variations of lift coefficients and drag coefficients with Mach numbers and angles of attack were demonstrated. The numerical results also indicate that the angle of canard deflection has significant influence on the flow field near tailfins.

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.

When a flat plate falls freely in periodic oscillating motion regime, unsteady fluid forces create additional lift force contributions due to the rotational behaviour. Computational fluid dynamics is used to simulate the free fall behaviour of a flat plate with aspect ratio β = 20 falling in two-dimensional flow with Reynolds number Re ≈ 10,000 and non-dimensional moment of inertia I* = 0.115. To validate the free fall trajectory obtained by computational fluid dynamics, video recordings are used. Based on the validated free fall computational fluid dynamics simulation, the instantaneous fluid forces and torques on the plate are obtained. The validated simulations show significant deviations in per-pendicular and tangential force coefficients at the same angle of attack depending on the trajectory history of the plate. At low angles of attack below 5 deg, the tangential force differs significantly. Oppositely, the difference in the perpendicular force is most pronounced at high angles of attack. During a free fall, the angle of attack is below 5 deg in 70 % of the time. Furthermore, the angle of attack is only above 45 deg in less than 5 % of the time. Therefore, effort must be put into a more detailed description of the tangential force component, in order to improve the existing modelling framework for non-spherical particles.

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.

Analytical prediction of trajectories of ice pieces after release in an airstream

This work provides an overview of aerodynamic data acquired in the European Transonic Windtunnel using an XRF-1 transport aircraft configuration both at cruise conditions and at the edges of the flight envelope. The goals and design of the wind tunnel test were described, highlighting the use of the cryogenic wind tunnel’s capability to isolate the effects of M_{infty }, {text{Re}}_{infty } and the dynamic pressure q/E. The resulting dataset includes an aerodynamic baseline characterization of the full span model with vertical and horizontal tailplanes and without engine nacelles. The effects of different inflow conditions were studied using data from continuous polars, evaluating the changes in aeroelastic deformation which are proportional to q/E and the influence of M_{infty } and {text{Re}}_{infty } on the shock position. Off-design data was analyzed at the lowest and highest measured Mach numbers of 0.84 and 0.90, respectively. Wing lower surface flow and underside shock motion was analyzed at negative angles of attack using {{c}_{text{p}}} distribution and unsteady pressure transducer fluctuation data, identifying significant upstream displacement of the shock close to the leading edge. Wing upper-side flow and the shock motion near buffet onset and beyond was analyzed using unsteady pressure data from point transducers and unsteady pressure-sensitive paint (PSP) measurements. Buffet occurs at lower angles of attack at high Mach number, and without clearly defined lift break. Spectral contents at the acquired data points in the buffet range suggest broadband fluctuations at Strouhal numbers between 0.2 and 0.6, which is consistent with recent literature. The spanwise shock propagation velocities were determined independently via analysis of unsteady PSP and pressure transducers to be in the range between u_{text{s}} / u_{infty } = 0.24 and 0.32, which is similarly in line with published datasets using other swept wing aircraft models.

A new method to quickly generate hypersonic waverider forebody configurations has been developed. The new method is an extension of the analysis method and power law method.and is called “Osculating analysis Method”. In this new method the constraint of creating only one attached wave shock in the forebody design by analysis or power law method is removed. More than one attached forebody shock waves can be created in the new method. For the purpose to integrate with ramjet or scramjet, a hypersonic forebody integrated with inlet was designed. The design objective is that the hypersonic forebody has three attached shock waves in the design flight condition. To better understand the forebody performance, three dimensional flow field calculation of the design waverider forebody (designed by osculating analysis method) integrated with hypersonic inlet and two comparative reference waverider forebodies(designed by analysis method and power law method, respectively) integrated with hypersonic inlet in the design and off design conditions were conducted. The computational results show that the performance of design waverider with three attached shock wave is great better than that of two comparative reference waveriders. Compared with two comparative reference models, the flow coefficient is increased by 10.38%, 8.2%; the recovery coefficient of total pressure is increased by 9.38%, 7%, respectively; the uniformity coefficient of inlet exit is increased by 8.8%,6.2%. Three waverider forebodies integrated with hypersonic inlet show better perforemance in high Mach number and attack angle than that in low mach number and negative attack angle. CFD predictions show that the design mach number of waverider forebody should be great than the real flight mach number and the design attack angle should be less than that of real flight condition. Test of the design waverider forebody integrated with hypersonic inlet was performed at a 500-mm-diameter hypersonic tunnel. Experimental results were obtained for a range of conditions, including the on-design Mach number and off-design Mach number. The experimental results were greatly agreed with those of simulation predictions. Vapor screen results for the model indicate three attached (or very nearly attached, within the limits of observation) shock waves along the entire leading edge. Simulation predictions generally agree with the measured pressure distribution.

A lift-driven vertical axis wind turbine (VAWT) generates peak power when it is rotating at high tip-speed ratios (TSR), at which time the blades encounter angles of attack (AOA) over a small range from zero to 30 degrees. However, its ability to self-start is dependent upon its performance at low TSRs, at which time the blades encounter a range of AOAs from zero to 180 degrees. A novel vented aerofoil is presented with the intention of improving the performance of a lift-driven VAWT at low TSRs without hampering the performance of the wind turbine at high TSRs. Computational fluid dynamics (CFD) simulation is used to predict the aerodynamic characteristics of a new vented aerofoil based on the well documented NACA0012 profile. Simulations are performed using the SST turbulence model. The results obtained show a reduction in the coefficient of tangential force (the force that generates torque on the wind turbine) at low AOAs (less than 90 degrees) of no more than 30%, while at high AOAs (more than 90 degrees) an improvement in the tangential force of over 100% is observed. Using a simple momentum based performance prediction model, these results suggest that this would lead to an increase in torque generation by a theoretical three-bladed VAWT of up to 20% at low TSRs and a minor reduction in coefficient of performance of up to 9% at TSR of 2 and closer to 1% at higher TSRs.

Modie cations to design of airfoils for use in unsteady transonic e ows are carried out by means of realistic geometry modeling and systematic removal of recompression shocks in the transonic region. A procedure for designing deformable airfoils for helicopter rotor blades is introduced. A baseline shock-free airfoil is created by combining systematic and optimization methods. Two analytical models of deformation are considered. Their ability to control dynamic stall is investigated numerically by using a two-dimensional, time-accurate, implicit Navier‐Stokes code. Results indicate the great potential for the controlled deformation of airfoils for helicopter applications. Dynamic stalling of airfoils at high angles of attack and the appearance of strong recompression shocks at low angles of attack and high Mach numbers are reduced signie cantly. In this way, the results presented will help in extending the design knowledge base for airfoils for helicopter rotors with advanced dynamic stall control.

This paper presents an evaluation of five different turbulence models by comparing the numerical data derived from these models using ANSYS Fluent with experimental data at a Reynolds number and a Mach number of 0.05 × 106 and 0.015 respectively based on the centerline chord of the airfoil for the flow over NACA 0012 and NACA 2412 airfoils. Moreover, the aim of the present study is to demonstrate the difference in aerodynamic characteristics of the airfoils in order to find aerodynamically more advantageous airfoil. It is concluded that Spalart-Allmaras model and k-ω SST model are capable of providing the most accurate prediction for lift coefficient at a low angle of attack for both airfoils. Standard k - ε model gives a slightly low value of lift coefficient at low angle of attack and slightly high value of lift coefficient at high angle of attack for both airfoils. k-ω SST model, Spalart-Allmaras model, Transition k-kL - ω model, and γ-Rⅇθ Transition SST model can predict drag coefficient reasonably at low angle of attack. At a high angle of attack, however, no turbulence model is able to give a satisfactory prediction for lift coefficient as well as drag coefficient, which implies that these models are unable to predict post-stall characteristics. NACA 2412 airfoil produces more lift coefficient than that of the NACA 0012 airfoil at all angles of attack. Moreover, the drag coefficient of NACA 2412 airfoil is less than that of the NACA 0012 airfoil, which implies that NACA 2412 airfoil exhibits better aerodynamic performance. The lift to drag coefficient ratio of NACA 2412 airfoil is also higher than that of the NACA 0012 airfoil indicating NACA 2412 airfoil to be more fuel economic.

In this paper the physics of the flow past a submerged vane at high angles of attack is presented. The data presented were collected during an experimental study of flow past vanes at various angles of attacks in a deformable-bed straight rectangular channel. The first few experimental runs were undertaken to establish the optimal angle of attack required to generate the strongest secondary circulation in the flow. At this optimal angle of attack, the flow features responsible for the transverse movement of bed sediment are clearly identified. The differences between the flow features observed near the vanes at both low and high angles of attack, are also outlined. These include, among others, two counterrotating vortices for vanes at high angles of attack as opposed to the widely acclaimed hypothesis of a single vortex for low angles of attack.

Investigations have been made adopting experiments and computations on an ogive-nosed slender body at different angles of attack and Reynolds number of 29,000 based on the model base diameter diameter. The results indicated an increase in the side force at large angles of attack, which is mainly due to the presence of asymmetric vortices in the leeward of the body. The inclusion of a rectangular cross-sectioned ring in the initial portion of the body reduced the side force at higher angles of attack. However, significant side force was experienced at lower angles of attack (30° < α < 40°). Use of a ring of 3% height was found to be suitable for reducing the side force at a higher angle of attack. From the results obtained it was observed that a ring if placed at a different axial location alters the flow field and changes the side force at higher angles of attack. Further studies indicated that placing of rings pair at an axial location of 3.5 and 4.5 times the base diameter reduced the side force to a very low value at all the angles of attack for the present shape of body and flow conditions.

Euler's equations, the shock wave relations, and the surface boundary condition are used to derive a hypersonic slender body theory applicable to angles of attack from zero to 90 degrees. The theory is not based on small disturbances, but it gives Van Dyke's theory for small angles of attack, Sychev's similarity for moderate angles of attack, and strip theory for large angles of attack. The new similarity parameters are Meo~, 0 cos a/~, and y (where ~=o cos a + sin a). They are shown to be a combination of Sychev's parameters and can be used to modify equations for pressure and force coefficients at small angles of attack for moderate and large angle of attack applications. B(X,y,Z) CA CA CD CD CN CN Cp Cp d D Db A n p,p S(x,y,Z) T Nomenclature scaled function for body geometry axial force coefficient, based on ref. area ::: 2Ld scaled axial force coefficient given by eq. (50) drag coefficient, based on ref. area ::: 2Ld scaled drag coefficient given by eq. (51) normal force coefficient based on ref. area ::: 2Ld scaled normal force coefficient given by eq. (19) pressure coefficient, eq. (47) ... scaled pressure coefficient given by eq. (47) base height of lower surface, see Fig. 1 parameter defined by eq. (2) parameter defmed by eq. (44) unit vectors in x, y, z directions, respectively similarity parameter, 0 cos a~ similarity parameter, Moo~ body length freestream Mach number unit vector normal (inner) to surface static pressure and scaled static pressure, eq. (25) scaled function for shock surface temperature * Professor of Mechanical and Aerospace Engineering, Associate Fellow AlAA. Copyright© 1992 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. 542 u,v,w u,v,w ~ Veo,VOQ x,y,z x,y,z a ~ y o ~ Cartesian velocity components in x,y,z directions, respectively scaled (nondimensional) velocity components freestream velocity vector and magnitude Cartesian coordinates, see Fig. 1 x/L, y/d, z/d, respectively angle of attack shockwave angle ratio of specific heats, 1.4 for air lower surface thickness ratio, d/L, see Fig. 1 --'> lower surface thickness ratio relative to V 00' see Fig. 1 and eq. (15) p,p density and scaled density, eq. (22) circumferential angle, =0 in windward plane SUbscripts c Cone b body s shock wave 00 freestream value Introduction Analysis of inviscid, hypersonic flow past slender bodies has been developed by Van Dyke l for small angles of attack and by Sychev2 for large angles of attack. The small angle of attack analysis is based on small disturbance theory and the similarity parameters are Moo 0, a/o, and y. At large angles of attack the Sychev similarity parameters are Moo sin a, 0 cot a, and y. Hemsch3 showed that the large angle of attack similarity parameters can be combined to yield the small angle of attack parameters when a is small, i.e. (MOQ sin a) (0 cot a) --'> Moo 0 and 0 cot a --'> o/a. Cox and Crabtree4 and Hayes and Probstein5 discuss hypersonic small disturbance theory and Sychev similarity. BarnweU6 showed that Sychev's analysis is applicable to all slender-body flows with crossflow Mach numbers greater than sonic and hence is not restricted to hypersonic values of the cross flow Mach number. In addition, he showed that Sychev similarity applied to a number of slender-body flows with subsonic crossflow Mach numbers, including incompressible flow. Hemsch3 demonstrated empirically that Sychev similarity holds for any value of the cross-flow Mach number if the axial flow component is 4

Multi-objective optimization of vortex generators on a transonic infinite wing is performed using computational fluid dynamics and a multi-objective genetic algorithm coupled with surrogate models. Vortex generator arrangements are defined by five design variables: height, length, incidence angle, chord location, and spacing. The objective functions are to maximize the lift-to-drag ratio at low angle of attack, to maximize lift coefficient at high angle of attack, and to the shift chordwise separation location downstream at high angle of attack. To evaluate these objective functions of each individual in the multi-objective genetic algorithm, the ordinary kriging surrogate model and the radial-basis-function/kriging hybrid surrogate model are employed because numerical analysis of the wing with vortex generators requires a large amount of computational time. Nondominated solutions are classified into five clusters with different aerodynamic characteristics. Comparison of the five clusters revealed that the balance among three objective functions is controlled mainly by vortex generator height, spacing, and their ratio. The solutions in each cluster have specific values of these three parameters, which identify the aerodynamic characteristics. In addition, appropriate values of design variables for generating the vortex most effectively are investigated.

Flow fields around an owl-like wing and aerodynamic characteristics at a chord Reynolds number of 23,000 are investigated using three-dimensional implicit large-eddy simulation. The cross sectional profile of the owl wing model named “owl-like wing” is constructed based on the owl wing at 40% of the span length from the root. It consists of flat upper surface, large camber, and thin geometry. Results show that at low angles of attack (α), separation, transition, and reattachment are observed in the instantaneous flow fields on the pressure side. The laminar separation bubbles can be seen in time- and span-averaged flow fields. It is likely that lift and drag generation is correlated with the location of separation points on the suction side. However, it has little influence on behavior of CL-α curve. On the other hand, at high angles of attack, the flow on the pressure side is fully attached. The flow on the suction side is similar to that of the pressure side at low angles of attack. It is found that unlike the case of the flow at the low angles of attack, the laminar separation bubble on the suction side affects the response of CL to variation of α. Furthermore, it is possible to decrease the drag and to increase the lift when the location of the laminar separation bubble is well organized by an appropriate airfoil surface geometry. Also, the deeply concaved lower surface contributes to lift enhancement. From those factors mentioned above, the owl-like wing gains higher lift-to-drag ratio comparing with conventional thin and thick symmetrical airfoils such as NACA0002 and NACA0012. Indeed, maximum lift-to-drag ratio of the owl-like wing is approximately 23 at the angle of attack of 6.0 degrees at Reynolds number of 23,000.