The application of a second generation low-order panel method - program 'Vsaero' - to powerplant installation studies

This study is a continuation of the work of Lee et al. (2024, “Aerodynamics and Flowfield of Distributed Electric Propulsion Tiltwing During Transition With Deflected Trailing-Edge Flap,” ASME J. Fluids Eng., 146(5), p. 051202) in which a large increase in lift and decrease in drag of a semiwing equipped with four four-bladed propellers at Re = 2.26 × 105 were reported. The force-balance measurements however contained both the thrust generated by the distributed electric propulsion (DEP) wing and the propeller slipstream-produced lift and drag forces. In the present experiment, the propeller slipstream-produced lift and drag were calculated through the surface pressures developed on the tilt DEP wing at multiple spanwise locations. By subtracting the propeller slipstream-produced lift and drag from the direct force-balance data, the thrust force generated by the DEP wing at each tilt angle and propeller rotation was obtained. The DEP wing-produced thrust was found to agree well with the surface pressure and force-balance measurements. The behavior of the boundary-layer flow developed on the DEP wing was also analyzed through the surface pressure measurements. It is anticipated that the present study will further advance our understanding of the aerodynamics of tiltwing electrical vertical take-off and landing(eVTOL) during transition maneuvering and its design.

A zonal method based on the iterative coupling between an in viscid panel method and a Navier-Stokes (NS) method has been developed to analyze the complex flowfield associated with the V/STOL flow environment in order to improve the prediction of aerodynamic characteristics. The low-order panel method, VSAERO, developed by Analytical Methods, Inc. is used to solve for the entire flow region, which includes the V/STOL configuration and jet boundary model. In two separate investigations presented here, ARC3D, a well-known compressible thin layer NS code developed at NASA Ames Research Center, and APPL, a parabolized NS code developed at Amtec Engineering, are used to solve for the details of the jet region subject to the boundary condition provided by VSAERO. Overlapping boundaries are used for the purpose of coupling. Comparisons with the available experimental data for a jet-in-crossflow and a V/STOL configuration with lift-jet in a subsonic freestream show qualitative agreement with experiment and improved correlations over those by inviscid computations alone.

A closed form for low-order panel methods

A BEM-isogeometric method for the ship wave-resistance problem

SummaryThree low-order panel methods developed for the analysis of supersonic flows have been used to make calculations for a number of test configurations, with the aim of establishing the range of problems for which this class of computational method yields adequate solutions. The programs investigated were the Woodward USSAERO/C program, the related NLRAERO program, and the British Aerospace Warton supersonic panel program. Results obtained using these programs have been evaluated against theoretical and experimental data for a number of test cases covering wing-alone, body-alone and wing-body geometries. It is concluded that low-order panel methods can provide adequate solutions for supersonic flows about wings, bodies and wing-body combinations, provided the assumptions implicit in the linearised potential flow model are not violated. Examples show that the prediction of lift and pitching moment curve slopes for quite complex configurations may be acceptable, but that the detailed pressure distributions are not always predicted satisfactorily. In particular, serious problems are encountered in calculating the flow about wings with rounded supersonic leading edges due to the linearised flow model which is used.

A low-order method is presented for aerodynamic prediction of wings operating at near-stall and post-stall flight conditions. The method is intended for use in design, modeling, and simulation. In this method, the flow separation due to stall is modeled in a vortex-lattice framework as an effective reduction in the camber, or "decambering." For each section of the wing, a parabolic decambering flap, hinged at the separation location of the section, is calculated through iteration to ensure that the lift and moment coefficients of the section match with the values from the two-dimensional viscous input curves for the effective angle of attack of the section. As an improvement from earlier low-order methods, this method also predicts the separation pattern on the wing. Results from the method, presented for unswept wings having various airfoils, aspect ratios, taper ratios, and small, quasi-steady roll rates, are shown to agree well with experimental results in the literature, and computational solutions obtained as part of the current work.

A B-spline based higher order panel method for analysis of steady flow around marine propellers

Flow around circular and noncircular cylinders by a low-order panel method

The predictions of fan performance on the pressure and flow rate characteristics of the axial flow fans using the conventional computational flow dynamics (CFD) approaches generally show a large deviation of about 20–30% from the experimental results. This article modifies the conventional CFD approaches by engaging the flow resistance of the fan test bench on the downstream area of the computational domain. The downstream pressure and flow rate are adjusted iteratively during the computational process. The results show that the prediction deviations using this method can be reduced to at most 3%, which is a great improvement compared to the conventional method. The inter-blade flows of three axial flow fans with different blade angles of attack are subsequently studied using this method. The results show that the flow patterns are drastically sensitive to the variation of the blade angle of attack, and the fan performances are closely related to the inter-blade flow behaviors. In the pre-stall regime, inappropriately designed blade angle of attack would cause the inter-blade flows in the region near the blade tip and the flows in the tip clearance region to present larger lateral and smaller axial velocity components with recirculation bubbles near the blade trailing edges. These flow behaviors cause the degradation of the fan performance. In the stall regime, boundary layer separation occurs to the suction surface of fan blades. Large recirculation bubbles appear near the trailing edges of blades and cause blockage effect against the axial flows. For the fans with inappropriately designed blade angle of attack, reverse flows can even be observed in the inter-blade passages with huge recirculation bubbles attaching to the trailing edges of fan blades.

Few studies have been devoted to effects of canard oscillations on the flowfield over the wing in supersonic flow and the associated time lags. In this article, an in-depth experimental study on a closely coupled canard-wing-body configuration in supersonic flow has been undertaken to mark the zone of influence of the oscillatory canard on the wing. Surface pressure measurements have been performed on the wing downstream of canard in both static and oscillatory cases at various wing angles of attack and canard deflection angles as well as the canard oscillation amplitudes and frequencies. The results show that the upwash and downwash flowfields due to an oscillatory canard periodically change the effective angle of attack seen by the wing and have different effects on various regions. In supersonic flow, the pressure rise due to the oblique shock waves causes the primary flow separation and the vortex break down onsets to occur at higher angles of attack, comparing to those in the lower speed regimes. According to the results, the mid-span region on the wing, which is directly in the shadow of the oscillating canard, receives more disturbances. During upstroke, the canard induces a downwash field to the front half of the wing, while in downstroke motion, the front regions of the wing experience upwash flow, giving rise to the effective angle of attack seen by the wing. Among various oscillation parameters, the pitching amplitude was observed to have a strong impact on the time lag in flowfield over the wing.

This study explores the design and development of a human-powered aircraft (HPA), leveraging modern engineering techniques, materials science, and advanced CAD/CAM tools. The project addresses key aspects of aircraft design, including the geometry of wings and tail, control and power transmission mechanisms, propeller selection, and material identification to achieve ultra-lightweight construction. The 3DExperience platform facilitated comprehensive model creation, simulation, and production process development, while XFLR5 was employed for aerodynamic profile analysis using the vortex lattice and panel methods. JavaProp aided in evaluating propeller thrust and power requirements. Computational fluid dynamics (CFD) simulations using the SST k-ω turbulence model provided critical insights into flow behavior. The design was found to be theoretically capable of flight, although challenges arose in selecting appropriate software for aerodynamic analysis, leading to the use of XFLR5 for early-stage design and the more advanced 3DExperience platform for final evaluations. Although structural strength analyses were not performed due to the complexity of composite materials, future work in this area could enhance the precision of component selection and aircraft mass estimation.

The purpose of the present paper is to determine numerically the effect of the shape of the nozzle of the air-cushion pad on the flow behavior near the nozzle. In the present paper, a new technique applied to the vortex method is proposed for case of calculations of various nozzle shapes : the fundamental shape of the nozzle is mapped conformally and the panel method is applied to various shapes of the nozzle exit in the mapping plane. The numerical calculation is compared with previous experimental results. Numerical results show that the velocity distribution at the nozzle exit agrees fairly well with experimental one and the global feature of streamlines is similar to experimental ones. The pressure distribution on the surface of strips oscillates remarkably due to the existence of discrete vortices near their surface, but it is pointed out that the flow is steady, in a sense of time average, except a few early stages and the pressure gap between the cushion chamber and the outside of the jet is shown.

Separation control experiments on a rectangular wing were carried out using dielectric barrier discharge plasma at subsonic speed for chord Reynolds numbers between 0,35 and 1·106 . Surface pressure measurements and flow visualization show that global flow separation on the wing can be mitigated or eliminated with the plasma actuators. The data were obtained for a wide range of angle of attack, flow speed, plasma excitation frequency and power. Various means of discharge influence on separated flow were investigated, including acoustic one

Wind tunnel tests were conduced to investigate roll characteristics of an SST configuration both with leading-edge flaps on the inboard and outboard wing and with trailing-edge flaps. Force measurements, surface pressure measurements and flow visualization tests were performed for the configuration with or without flap deflection. The test results indicated that rolling moments are stable for the models both with and without flap deflection at low incidence angles. At high incidence angles, unstable rolling moment was observed at limited roll angle for the model without flap deflection. Drastic change of rolling moment was observed at higher incidence angles when each flaps deflection.

The added mass coefficients for two and three-dimensional submerged bodies were calculated using a low-order panel code. The source and dipole strengths, and the panel surface area for each panel, were used to compute the integrals needed for added mass in all six degrees of motions. Several applications of this method were used in comparing the results with the theoretical, when available, experimental or other numerical results. The method was found to be successful in predicting the added mass coefficients using relatively low numbers of panels.