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Thermal stresses, which may result from transient heating, can cause changes in the effective stiffness of wing structures. Some effects of this change in stiffness were investigated experimentally by radiantly heating three types of simple wing structures: a uniform plate, a solid double-wedge section, and a circular-arc multiweb-wing section. Changes in stiffness were determined by measuring the changes in natural frequency of vibration during transient heating. Some comparisons are made between theoretical calculations and the measured data.
The main object of these wings was to obtain a high lift without increasing the total drag, and the idea was conceived that, by using a deep high-lift section for the top wing and a medium section for the bottom wing, the mutual interference between the two would be such as to give greater efficiency of the combination as compared with two wings of equal section.
Approximation method for calculating spanwise distribution of wind tunnel boundary interference on wing lift in rectangular perforated wall
A method is presented for calculating wing characteristics by lifting-line theory using nonlinear section lift data. Material from various sources is combined with some original work into the single complete method described. Multhopp's systems of multipliers are employed to obtain the induced angle of attack directly from the spanwise lift distribution. Equations are developed for obtaining these multipliers for any even number of spanwise stations, and values are tabulated for 10 stations along the semispan for asymmetrical, symmetrical, and antisymmetrical lift distributions. In order to minimize the computing time and to illustrate the procedures involved, simplified computing forms containing detailed examples are given for symmetrical lift distributions. Similar forms for asymmetrical and antisymmetrical lift distributions, although not shown, can be readily constructed in the same manner as those given. The adaptation of the method for use with linear section lift data is also illustrated. The adaptation has been found to require less computing time than most existing methods.
Some results are described for a lifting rectangular wing centrally located on a circular-cylindrical body. This simple configuration has been utilized in order to assess the merits of a mapping technique for wing-body configurations. The procedure employed makes use of a coordinate transformation to simplify specification of the surface boundary condition in the computation of the flow about the wing. The method can be extended to incorporate wing sweep, finite length body of noncircular cross section, and arbitrary wing placement; however, these extensions involve a considerable increase in complexity of the problem.
Unsteady transonic flow calculations are presented for wing-fuselage configurations. Calculations are performed by extending the XTRAN3S unsteady transonic small-disturbance code to allow the treatment of a fuselage. Details of the XTRAN3S fuselage modeling are discussed in the context of the small-disturbance equation. Transonic calculations are presented for three wing-fuselage configurations with leading edge sweep angles ranging from 0 deg to 46.76 deg. Simple bending and torsion modal oscillations of the wing are calculated. Sectional lift and moment coefficients for the wing-alone and wing-fuselage cases are compared and the effects of fuselage aerodynamic interference on the unsteady wing loading are revealed. Tabulated generalized aerodynamic forces used in flutter analyses, indicate small changes in the real in-phase component and as much as a 30% change in the imaginary component when the fuselage is included in the calculation. These changes result in a 2 to 5% increase in total magnitude and a several degree increase in phase.
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Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional icing simulation capabilities. An understanding of ice-shape geometric fidelity and Reynolds and Mach number effects on the iced-wing aerodynamics is needed to guide the development and validation of ice-accretion simulation tools. To this end, wind-tunnel testing and computational flow simulations were carried out for an 8.9 percent-scale semispan wing based upon the Common Research Model airplane configuration. The wind-tunnel testing was conducted at the Wichita State University 7 by 10 ft Beech wind tunnel from Reynolds numbers of 0.8×10(exp 6) to 2.4×10(exp 6) and corresponding Mach numbers of 0.09 to 0.27. This paper presents the results of initial studies investigating the model mounting configuration, clean-wing aerodynamics and effects of artificial ice roughness. Four different model mounting configurations were considered and a circular splitter plate combined with a streamlined shroud was selected as the baseline geometry for the remainder of the experiments and computational simulations. A detailed study of the clean-wing aerodynamics and stall characteristics was made. In all cases, the flow over the outboard sections of the wing separated as the wing stalled with the inboard sections near the root maintaining attached flow. Computational flow simulations were carried out with the ONERA elsA software that solves the compressible, threedimensional RANS equations. The computations were carried out in either fully turbulent mode or with natural transition. Better agreement between the experimental and computational results was obtained when considering computations with free transition compared to turbulent solutions. These results indicate that experimental evolution of the clean wing performance coefficients were due to the effect of three-dimensional transition location and that this must be taken into account for future data analysis. This research also confirmed that artificial ice roughness created with rapid-prototype manufacturing methods can generate aerodynamic performance effects comparable to grit roughness of equivalent size when proper care is exercised in design and installation. The conclusions of this combined experimental and computational study contributed directly to the successful implementation of follow-on test campaigns with numerous artificial ice-shape configurations for this 8.9 percent scale model.
Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional icing simulation capabilities. An understanding of ice-shape geometric fidelity and Reynolds and Mach number effects on the iced-wing aerodynamics is needed to guide the development and validation of ice-accretion simulation tools. To this end, wind-tunnel testing and computational flow simulations were carried out for an 8.9%-scale semispan wing based upon the Common Research Model airplane configuration. The wind-tunnel testing was conducted at the Wichita State University 7 ft x 10 ft Beech wind tunnel from Reynolds numbers of 0.8×10(exp 6) to 2.4×10(exp 6) and corresponding Mach numbers of 0.09 to 0.27. This paper presents the results of initial studies investigating the model mounting configuration, clean-wing aerodynamics and effects of artificial ice roughness. Four different model mounting configurations were considered and a circular splitter plate combined with a streamlined shroud was selected as the baseline geometry for the remainder of the experiments and computational simulations. A detailed study of the clean-wing aerodynamics and stall characteristics was made. In all cases, the flow over the outboard sections of the wing separated as the wing stalled with the inboard sections near the root maintaining attached flow. Computational flow simulations were carried out with the ONERA elsA software that solves the compressible, three-dimensional RANS equations. The computations were carried out in either fully turbulent mode or with natural transition. Better agreement between the experimental and computational results was obtained when considering computations with free transition compared to turbulent solutions. These results indicate that experimental evolution of the clean wing performance coefficients were due to the effect of three-dimensional transition location and that this must be taken into account for future data analysis. This research also confirmed that artificial ice roughness created with rapid-prototype manufacturing methods can generate aerodynamic performance effects comparable to grit roughness of equivalent size when proper care is exercised in design and installation. The conclusions of this combined experimental and computational study contributed directly to the successful implementation of follow-on test campaigns with numerous artificial ice-shape configurations for this 8.9% scale model.
A computer code, XSECT, has been developed to generate fuselage cross sections from a given area distribution and wing definition. The cross sections are generated to match the wing definition while conforming to the area requirement. An iterative procedure is used to generate each cross section. Fuselage area balancing may be included in this procedure if desired. The code is intended as an aid for engineers who must first design a wing under certain aerodynamic constraints and then design a fuselage for the wing such that the contraints remain satisfied. This report contains the information necessary for accessing and executing the code, which is written in FORTRAN to execute on the Cyber 170 series computers (NOS operating system) and produces graphical output for a Tektronix 4014 CRT. The LRC graphics software is used in combination with the interface between this software and the PLOT 10 software.
Aerodynamic characteristics of a general aviation wing equipped with NACA 65 sub 2-415, NASA GA(W)-1, and NASA GA(PC)-1 airfoil sections were examined. The NASA GA(W)-1 wing was equipped with plain, split, and slotted partial- and full-span flaps and ailerons. The NASA GA(PC)-1 wing was equipped with plain, partial- and full-span flaps. Experimental chordwise static-pressure distribution and wake drag measurements were obtained for the NASA GA(PC)-1 wing at the 22.5-percent spanwise station. Comparisons were made between the three wing configurations to evaluate the wing performance, stall, and maximum lift capabilities. The results of this investigation indicated that the NASA GA(W)-1 wing had a higher maximum lift capability and almost equivalent drag values compared with both the NACA 65 sub 2-415 and NASA GA(PC)-1 wings. The NASA GA(W)-1 had a maximum lift coefficient of 1.32 with 0 deg flap deflection, and 1.78 with 41.6 deg deflection of the partial-span slotted flap. The effectiveness of the NASA GA(W)-1 plain and slotted ailerons with differential deflections were equivalent. The NASA GA(PC)-1 wing with full-span flaps deflected 0 deg for the design climb configuration showed improved lift and drag performance over the cruise flap setting of -10 deg.
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The problem of autorotation has been attacked by various authors, in particular Fuchs and Schmidt who applied the so-called strip method which is based upon the assumption that the forces and moments per unit length acting in each section of the wing are equal to those on an infinite cylindrical wing of equal section in an air flow of intensity and direction resulting from the apparent relative motion of this section with respect to the surrounding air. In other words they disregarded the induced velocities, which, however, are of such importance that their omission is bound to result in appreciable errors, as we attempt to prove in this report.
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This report gives the results of an investigation conducted in the NACA full-scale wind tunnel on a small parasol monoplane equipped with three different split trailing-edge wing flaps. The object of the investigation was to determine and correlate data on the characteristics of the airplane and flaps as affected by variation in flap chord, flap deflection, and flap location along the wing chord. The results give the lift, the drag, and the pitching moment characteristics of the airplane, and the flap forces and moments, the pressure distribution over the flaps and wing at one section, and the downwash characteristics of the flap and wing combinations.
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