Downwash Survey Behind Two Low-aspect-ratio Variable-incidence Wings in Combination with Three Different Size Fuselages at a Mach Number of 0.25
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Hopkins, Edward J↗
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Low-subsonic Investigation to Determine the Chordwise Pressure Distribution and Effectiveness of Spoilers on a Thin, Low-aspect-ratio, Unswept, Untapered, Semispan Wing and on the Wing with Leading- and Trailing-edge Flaps
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Croom, Delwin R↗
Analysis of Pressure Data Obtained at Transonic Speeds on a Thin Low-aspect-ratio Cambered Delta Wing-body Combination
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Mugler, John P , Jr↗
Wind-tunnel investigation at subsonic and supersonic speeds of a model of a tailless fighter airplane employing a low-aspect-ratio swept-back wing : effects of external fuel tanks and rocket packets on the drag characteristics
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Smith, Willard G↗
Zero lift-drag at mach 1.42, 1.83, and 2.21 of a series of wings with variations of thickness ratio and chord
Zero lift-drag at supersonic speeds of wings with low thickness ratios
LOW ASPECT RATIO WING↗
Investigation at low speeds of the effect of aspect ratio and sweep on rolling stability derivatives of untapered wings
A low scale wind tunnel investigation was conducted in rolling flow to determine the effects of aspect ratio and sweep (when varied independently) on the rolling stability derivatives for a series of untapered wings. Test results indicate that when the aspect ratio was held constant, an increase in the sweepback angle caused a significant reduction in the damping in roll at low lift coefficients for only the higher aspect ratios that were tested. This result was in agreement with available swept wing theory which indicated no effect of sweep for aspect ratios near zero. The result of the linear theory that the damping in roll is independent of lift coefficient and that the yawing moment and lateral force due to rolling are directly proportional to the lift coefficient was found to be valid for only a very limited lift coefficient range when the wings were highly swept. For such wings, the damping was found to increase in magnitude and the yawing moment due to rolling, to change from negative to positive at moderate lift coefficients. The effect of wing tip suction, not acounted for by present theory, was found to be very important with regard to the yawing moment due to rolling, particularly for low aspect ratio swept wings. An empirical means of correcting present theory for the effect of tip suction is suggested.
Goodman, Alex↗
Stability derivatives of triangular wings at supersonic speeds
The analysis of the stability derivatives of low-aspect-ratio triangular wings at subsonic and supersonic speeds, given in NACA TN no. 1423, is extended to apply to triangular wings having large vertex angles and traveling at supersonic speeds. The lift, rolling moment due to sideslip, and damping in roll and pitch for this more general case have been treated elsewhere on the basis of the theory of small disturbances. The surface potentials for angle of attack and rolling taken therefrom are used to obtain the several side-force and yawing-moment derivatives that depend on leading-edge suction, and a tentative value for the rolling moment due to yawing. The lift and moment due to downward acceleration are obtained on the basis of an unpublished unsteady-flow solution. All the known stability derivatives of the triangular wing at supersonic speeds, regardless of source, are summarized for convenience and presented with respect to both body axes and stability axes. The results are limited to Mach numbers for which the triangular wing is contained within the Mach cone for its vertex. The spanwise variation of Mach number in the case of yawing is neglected, although the effect must be of importance.
Ribner, Herbert S↗
Effect of drooped-nose flaps on the experimental force and moment characteristics of an oblique wing
Six-component experimental force and moment data are presented for a low aspect ratio, oblique wing equipped with drooped-nose flaps and mounted on top of a body of revolution. These flaps were investigated on the downstream wing panel with the nose drooped 5 deg, 10 deg, 20 deg, and 30 deg, and on both wing panels with the nose drooped 30 deg. It was to determine if such flaps would make the moment curves more linear by controlling the flow separation on the downstream wing panel at high lift coefficients. The wing was elliptical in planform and had an aspect ratio of 6.0 (based on the unswept wing span). The wing was tested at sweep angles of 45 deg and 50 deg throughout the Mach number range from 0.25 to 0.95. The drooped-nose flaps alone were not effective in making the moment curves more linear; however, a previous study showed that Kruger nose flaps improved the linearity of the moment curves when the Kruger flaps were used on only the downstream wing panel equipped with drooped-nose flaps deflected 5 deg.
Hopkins, E. J.↗
A Survey of Factors Affecting Blunt Leading-Edge Separation for Swept and Semi-Slender Wings
A survey is presented of factors affecting blunt leading-edge separation for swept and semi-slender wings. This class of separation often results in the onset and progression of separation-induced vortical flow over a slender or semi-slender wing. The term semi-slender is used to distinguish wings with moderate sweeps and aspect ratios from the more traditional highly-swept, low-aspect-ratio slender wing. Emphasis is divided between a selection of results obtained through literature survey a section of results from some recent research projects primarily being coordinated through NATO s Research and Technology Organization (RTO). An aircraft context to these studies is included.
Luckring, James M.↗
Analytical observations on the aerodynamics of a delta wing with leading edge flaps
The effect of a leading edge flap on the aerodynamics of a low aspect ratio delta wing is studied analytically. The separated flow field about the wing is represented by a simple vortex model composed of a conical straight vortex sheet and a concentrated vortex. The analysis is carried out in the cross flow plane by mapping the wing trace, by means of the Schwarz-Christoffel transformation into the real axis of the transformed plane. Particular attention is given to the influence of the angle of attack and flap deflection angle on lift and drag forces. Both lift and drag decrease with flap deflection, while the lift-to-drag ratioe increases. A simple coordinate transformation is used to obtain a closed form expression for the lift-to-drag ratio as a function of flap deflection. The main effect of leading edge flap deflection is a partial suppression of the separated flow on the leeside of the wing. Qualitative comparison with experiments is presented, showing agreement in the general trends.
Oh, S.↗
Test Cases for a Rectangular Supercritical Wing Undergoing Pitching Oscillations
Steady and unsteady measured pressures for a Rectangular Supercritical Wing (RSW) undergoing pitching oscillations have been presented. From the several hundred compiled data points, 27 static and 36 pitching oscillation cases have been proposed for computational Test Cases to illustrate the trends with Mach number, reduced frequency, and angle of attack. The wing was designed to be a simple configuration for Computational Fluid Dynamics (CFD) comparisons. The wing had an unswept rectangular planform plus a tip of revolution, a panel aspect ratio of 2.0, a twelve per cent thick supercritical airfoil section, and no twist. The model was tested over a wide range of Mach numbers, from 0.27 to 0.90, corresponding to low subsonic flows up to strong transonic flows. The higher Mach numbers are well beyond the design Mach number such as might be required for flutter verification beyond cruise conditions. The pitching oscillations covered a broad range of reduced frequencies. Some early calculations for this wing are given for lifting pressure as calculated from a linear lifting surface program and from a transonic small perturbation program. The unsteady results were given primarily for a mild transonic condition at M = 0.70. For these cases the agreement with the data was only fair, possibly resulting from the omission of viscous effects. Supercritical airfoil sections are known to be sensitive to viscous effects (for example, one case cited). Calculations using a higher level code with the full potential equations have been presented for one of the same cases, and with the Euler equations. The agreement around the leading edge was improved, but overall the agreement was not completely satisfactory. Typically for low-aspect-ratio rectangular wings, transonic shock waves on the wing tend to sweep forward from root to tip such that there are strong three-dimensional effects. It might also be noted that for most of the test, the model was tested with free transition, but a few points were taken with an added transition strip for comparison. Some unpublished results of a rigid wing of the same airfoil and planform that was tested on the pitch and plunge apparatus mount system (PAPA) showed effects of the lower surface transition Strip on flutter at the lower subsonic Mach numbers. Significant effects of a transition strip were also obtained on a wing with a thicker supercritical section on the PAPA mount system. Both of these flutter tests on the PAPA resulted in very low reduced frequencies that may be a factor in this influence of the transition strip. However, these results indicate that correlation studies for RSW may require some attention to the estimation of transition location to accurately treat viscous effects. In this report several Test Cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flow effects. An overview of the model and tests is given and the standard formulary for these data is listed. Sample data points are presented in both tabular and graphical form. A complete tabulation and plotting of all the Test Cases is given. Only the static pressures and the real and imaginary parts of the first harmonic of the unsteady pressures are available. All the data for the test are available in electronic file form. The Test Cases are also available as separate electronic files.
Bennett, Robert M.↗
High-Speed Wind-Tunnel Tests of a 1/16-Scale Model of the D-558 Research Airplane Lift and Drag Characteristics of the D-558-1 and Various Wing and Tail Configurations
Tests were made in the Langley 8-foot high-speed tunnel to investigate the aerodynamic characteristics of the D-558-1 airplane and various wing and tail configurations on the D-558-1 fuselage. The various wing and tail configurations were tested to determine the aerodynamic effects of aspect ratio and sweep for suitable use on the second phase of the D-558 project (D-558-2). The tests were conducted through a speed range from a Mach number of 0.40 to approximately 0.94.This part of the investigation includes the lift and drag results available for the configurations tested at this rate. The D-558-1 results indicated that the lift force break would occur at a Mach number of 0.85 with some reduction in lift at speeds above this Mach number. Tests indicated that the airplane will have satisfactory lift and drag characteristics up to and including its design Mach number of 0.85. The 35deg sweptback, 35deg swept-forward, and low-aspect-ratio (2.0) wing configurations all showed pronounced improvements in maintaining lift throughout the Mach number range tested and in increasing the critical speeds above the D-558-1 value &itical to critical Mach numbers on the order of 0.9. Insofar as lift and drag characteristics are concerned level flight at speeds approaching the velocity of sound appears practical if swept or low-aspect-ratio configurations similar to those tested are used.
Wright, John D.↗
The Aerodynamic Forces on Slender Plane- and Cruciform-Wing and Body Combinations
The load distribution, forces, and moments are calculated theoretically for inclined slender wing-body combinations consisting of a slender body of revolution and either a plane or cruciform arrangement of low-aspect-ratio pointed wings. The results are applicable at subsonic and transonic speeds, and at supersonic speeds, provided the entire wing-body combination lies near the center of the Mach cone.
Spreiter, John R↗
A LEX blowing technique for post-stall lateral control of trapezoidal wings
An exploratory low-speed wind-tunnel investigation of a pneumatic technique for LEX-induced lift augmetation on a low-aspect-ratio trapezoidal wing near C sub L, MAX is reported. The technique involved lateral sheet ejection from LEX edge-slots, with emphasis on nonsymmetrical blowing to obtain incremental roll control at high angles of attack when the aileron power is declining. Preliminary experiments on a pressure instrumented semispan wing model followed by force/moment measurements on a full-span LEX-wing configuration confirmed the effectiveness of LEX blowing for lateral control improvement up to an angle of attack of 45 deg and with a momentum coefficient of less than 02.
Rao, Dhanvada M.↗
Applications of oblique-wing technology - An overview
A summary of oblique-wing technology programs, including system studies and experimental activities, is presented. Oblique-wing civil transports were studied and compared with conventional configurations at design cruise Mach numbers of 0.95 and 1.2. Studies of military configurations employing oblique-wing technology considered aircraft concepts, remotely piloted vehicles, and a cruise missile. Experimental activities included wind-tunnel tests of high and low aspect ratio oblique wings; flight tests are planned using unmanned and manned vehicles. The research program suggests that oblique-wing technology offers improved performance over conventional concepts for aircraft operating in the transonic speed range.
Nelms, W. P., Jr.↗
Aeroacoustic Computations of a Generic Low Boom Concept in Landing Configuration: Part 1 - Aerodynamic Simulations
Computational results are presented for a 15%-scale, full-span, generic low-boom concept aircraft model simulated as installed in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel and in free-air conditions. The simulations were conducted with the lattice Boltzmann solver PowerFLOW® to capture the time-accurate characteristics of the flow. The aerodynamic behavior of the aircraft was investigated with various combinations of flap and landing gear deployments. To replicate the flow speed used during tests of the model, a Mach number of approximately 0.20 was chosen while varying the angle of attack through a limited range. Analyses were performed on local flow quantities, global forces, as well as flow distributions between configurations. Aerodynamic quantities were shown to be sensitive to mesh resolution levels, especially with flap and landing gear systems retracted. When comparing wind tunnel to free-air simulations, the aerodynamic quantities sampled were found to be in good agreement, suggesting that tunnel blockage and model support effects did not adversely impact the aerodynamic behavior of the model during the wind tunnel test. The results of this study improve our understanding of the highly complex, vortical flow generated by configurations with swept, low aspect-ratio wings at speeds encountered during landing operations.
low boom↗
Wind-Tunnel Investigation at Subsonic and Supersonic Speeds of the Static and Dynamic Stability Derivatives of an Airplane Model with an Unswept Wing and a High Horizontal Tail
Results are presented of a wind-tunnel investigation to evaluate the static and dynamic stability derivatives of a model with a low-aspect-ratio unswept wing and a high horizontal tail. In addition to results for the complete model, results were also obtained of the body alone, body and wing, and body and tail. Data were obtained in the Mach number range from 0.65 to 2.2, at a Reynolds number of 2 million based on the wing mean aerodynamic chord. The angle-of-attack range for most of the data was -11.5 deg to 18 deg. A limited amount of data was obtained with fixed transition. A correspondence between the damping in pitch and the static stability, previously noted in other investigations, was also observed in the present results. The effect observed was that a decrease (or increase) in the static stability was accompanied by an increase (or decrease) in the damping in pitch. A similar correspondence was observed between the damping in yaw and the static-directional stability. Results from similar tests of the same model configuration in two other facilities over different speed ranges are presented for comparison. It was found that most of the results from the three investigations correlated reasonably well. Estimates of the rotary derivatives were made using available procedures. Comparison with the experimental results indicates the need for development of more precise estimation procedures.
Lessing, Henry C.↗
Recent Results Pertaining to the Application of the "Area Rule"
This paper is concerned primarily with the application of the "area rule" to the interpretation and improvement of the drag-rise characteristics of wing-body combinations at transonic and moderate supersonic speeds. Consideration of the general physical nature of the flow at transonic speeds, together with comparisons of the flow fields and drag-rise characteristics for wing-body combinations and bodies of revolution has led to the conclusion that near the speed of sound the drag rise for a thin low-aspect-ratio wing-body combination is primarily dependent on the axial distribution of cross-sectional area normal to the airstream (ref. 1). (The drag rise, sometimes referred to as pressure drag, is the difference between the drag level near the speed of sound and the drag level at subsonic speeds where the drag is due primarily to skin friction.) In order to illustrate the concept, figure 1 shows a wing-body combination and a body of revolution. A typical cross section normal to the airstream for the wing-body combination is shown at AA. The cross-sectional area of the wing is wrapped around the body of revolution so that the body has the same cross-sectional area at BB. All the other cross-sectional areas of the body of revolution are the same as those for the wing-body combination at the same axial stations. On the basis of the conclusion just stated, the drag rise for this body of revolution should be similar to that for the wing-body combination.
Whitcomb, Richard T.↗