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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A comparison at Mach numbers up to 0.92 of the calculated and experimental downwash and wake characteristics at various horizontal tail heights behind a wing with 45 degree of sweepback
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Stephenson, Jack D↗
An Experimental Investigation of the Unsteady Lift Induced on a Wing in the Downwash Field of an Oscillating Canard Control Surface
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Reese, David E , Jr↗
Some Effects of Sweep and Thickness on the Experimental Downwash Characteristics at Transonic Speeds of a Series of Highly Tapered Wings with an Aspect Ratio of 3 : Transonic Bump Method
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Few, Albert G , Jr↗
An Analysis of Pressure Studies and Experimental and Theoretical Downwash and Sidewash Behind Five Pointed-Tip Wings at Supersonic Speeds
Flow-angle and pressure surveys behind five, thin, pointed-tip wings of varying plan form have been made at Mach numbers 1.62 and 2.41. Schlieren studies at a Mach number 1.93 for the same five plan-form wings were made to illustrate the behavior of the vortex sheet. The surveys were conducted at 1.5, 3, and 4 root chords behind three triangular wings of 50 degree, 63 degree, and 72 degree leading-edge sweep angle, and behind the 50 degree triangular wing reversed. The flow behind a pointed-tip wing having a sweptback leading edge and a sweptforward trailing edge (both 50 degrees) was also surveyed. Experiment and one of the theoretical methods are compared for the reversed triangular wing and the pointed-tip wing with the 50 degree sweptback leading edge and sweptforward trailing edge.
Boatright, William B↗
Downwash measurements behind wings with detached float
This investigation, which was made in the small wind tunnel having a diameter of 1.2 m (3.94 feet), embraced three wing models, behind which, at various angles of attack between 0 and 60 degrees, the static pressure and the total pressure along vertical lines (perpendicular to the direction of the undisturbed wind and to the wing span) were measured. The location of these vertical lines are indicated in Figure 1. Moreover, the wing polars were determined by the customary three-component measurements. For testing the pressure field, a Pitot tube and a static probe, both of 2 mm (0.08 in.) in diameter, were mounted 40 mm (1.57 in.) apart on the end of a shaft 1 m (39.37 in.) long.
Petersohn, E↗
An Investigation of the Effect of Downwash from a VTOL Aircraft and a Helicopter in the Ground Environment
Dynamic-pressure measurement, in ground effect, have been obtained about a single-rotor helicopter and a dual-propeller VTOL (Vertical Take-Off and Landing) aircraft. The results indicate that the slipstream dynamic pressure along the ground, some distance from the center of rotation, is not a function of disk loading but merely a function of the gross weight or thrust of the aircraft. Furthermore, for a given gross weight the thickness of this outward flowing sheet of air is less for a small-diameter propeller (higher disk loading propeller). The variation of the dynamic-pressure flow field for single and dual propellers or rotors is significantly different in the plane of symmetry between the two rotors than in a direction normal to this plane. The interaction of the two flows produces a region of upflow in this plane where the fuselage is located, and the decay of the maximum dynamic pressure with distance ahead of the fuselage is slower.
O'Bryan, Thomas C.↗
Lifting and transport of Martian dust by the Ingenuity helicopter rotor downwash as observed by high-speed imaging from the Perseverance rover
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Mark T Lemmon↗
Computation of aerodynamic interference effects on oscillating airfoils with controls in ventilated subsonic wind tunnels
Lift interference effects are discussed based on Bland's (1968) integral equation. A mathematical existence theory is utilized for which convergence of the numerical method has been proved for general (square-integrable) downwashes. Airloads are computed using orthogonal airfoil polynomial pairs in conjunction with a collocation method which is numerically equivalent to Galerkin's method and complex least squares. Convergence exhibits exponentially decreasing error with the number n of collocation points for smooth downwashes, whereas errors are proportional to 1/n for discontinuous downwashes. The latter can be reduced to 1/n to the m+1 power with mth-order Richardson extrapolation (by using m = 2, hundredfold error reductions were obtained with only a 13% increase of computer time). Numerical results are presented showing acoustic resonance, as well as the effect of Mach number, ventilation, height-to-chord ratio, and mode shape on wind-tunnel interference. Excellent agreement with experiment is obtained in steady flow, and good agreement is obtained for unsteady flow.
Fromme, J. A.↗
Math modeling for helicopter simulation of low speed, low altitude and steeply descending flight
A math model was formulated to represent some of the aerodynamic effects of low speed, low altitude, and steeply descending flight. The formulation is intended to be consistent with the single rotor real time simulation model at NASA Ames Research Center. The effect of low speed, low altitude flight on main rotor downwash was obtained by assuming a uniform plus first harmonic inflow model and then by using wind tunnel data in the form of hub loads to solve for the inflow coefficients. The result was a set of tables for steady and first harmonic inflow coefficients as functions of ground proximity, angle of attack, and airspeed. The aerodynamics associated with steep descending flight in the vortex ring state were modeled by replacing the steady induced downwash derived from momentum theory with an experimentally derived value and by including a thrust fluctuations effect due to vortex shedding. Tables of the induced downwash and the magnitude of the thrust fluctuations were created as functions of angle of attack and airspeed.
Sheridan, P. F.↗
Modeling of Longitudinal Unsteady Aerodynamics of a Wing-Tail Combination
Aerodynamic equations for the longitudinal motion of an aircraft with a horizontal tail were developed. In this development emphasis was given on obtaining model structure suitable for model identification from experimental data. The resulting aerodynamic models included unsteady effects in the form of linear indicial functions. These functions represented responses in the lift on the wing and tail alone, and interference between those two lifting surfaces. The effect of the wing on the tail was formulated for two different expressions concerning the downwash angle at the tail. The first expression used the Cowley-Glauert approximation known-as "lag-in-downwash," the second took into account growth of the wing circulation and delay in the development of the lift on the tail. Both approaches were demonstrated in two examples using the geometry of a fighter aircraft and a large transport. It was shown that the differences in the two downwash formulations would increase for an aircraft with long tail arm performing low-speed, rapid maneuvers.
Klein, Vladislav↗
Wake Vortex Effects Between Urban Air Mobility Aircraft
This study models the wake-induced downwash of an Urban Air Mobility (UAM) aircraft and the wake’s effects on a trailing aircraft’s dynamic stability when intersecting the wake trail. The Proctor wake vortex model is used to compute the wake-induced downwash over a 0.8 × 0.8 nautical mile flight area. The model is complemented by core radius growth, Proctor single phase wake decay, wake propagation, and Lamb wake aging. The wake data is then stored and loaded to the trailing aircraft as an atmospheric condition to affect the dynamics. Preliminary simulations include modeling the downwash from multiple UAM aircraft over the flight area and determining the normalized wake reduction at the ICAO 5 nautical mile separation criteria for a LIGHT aircraft trailing a MEDIUM aircraft.
Wake Vortex↗
Longitudinal Aerodynamic Characteristics to Large Angles of Attack of a Cruciform Missile Configuration at a Mach Number of 2
The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing-tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift-interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance between the wing trailing vortex wake and the tail surfaces and thus was a function of angle of attack, angle of roll, and wing-tail interdigitation. Although the configuration at zero roll with the wing and tail in line exhibited the least center-of-pressure travel, the configuration with the wing and tail interdigitated had the least change in wing-tail interference over the angle-of-attack range. The lift effectiveness of the variable-incidence wing was reduced by more than 70 percent as a result of an increase in the combined angle of attack and wing incidence from 0 deg to 40 deg. The wing-tail interference (effective downwash at the tail) due to wing deflection was nearly zero as a result of a region of negative vorticity shed from the inboard portion of the wing. The lift characteristics of the configuration and its components were satisfactorily predicted by the calculated results, but the pitching moments at large angles of attack were not because of the influence of factors for which no adequate theory is available, such as the variation of the crossflow drag coefficient along the body and the effect of the wing downwash field on the afterbody loading.
Spahr, J. R.↗
Prediction of unsteady aerodynamic loadings caused by trailing edge control surface motions in subsonic compressible flow: Analysis and results
A theoretical analysis and a computer program have been developed for the prediction of unsteady lifting surface loadings caused by motions of trailing edge control surfaces having sealed gaps. The final form of the downwash integral equation has been formulated by isolating the singularities from the non-singular terms and establishing a preferred solution process to remove and evaluate the downwash discontinuities in a systematic manner. Comparisons of theoretical and experimental pressure data are made for several control surface configurations. The comparisons indicate that reasonably accurate theoretical pressure distributions and generalized forces may be obtained for a wide variety of control surface configurations. Spanwise symmetry or antisymmetry of motion, and up to four control surfaces on each half span can be accommodated.
Rowe, W. S.↗
Research on unsteady transonic flow theory
A two-dimensional theory is considered for the unsteady flow disturbances caused by aeroelastic deformations of a thick wing at high subsonic freestream Mach numbers, having a single, internally embedded supercritical (locally supersonic) steady flow region adjacent to the low pressure side of the wing. The theory develops a matrix of unsteady aerodynamic influence coefficients (AICs) suitable as a strip theory for aeroelastic analysis of large aspect ratio thick wings of moderate sweep, typical of a wide class of current and future aircraft. The theory derives the linearized unsteady flow solutions separately for both the subcritical and supercritical regions. These solutions are coupled together to give the requisite (wing pressure-downwash) AICs by the intermediate step of defining flow disturbances on the sonic line, and at the shock wave; these intermediate quantities are then algebraically eliminated by expressing them in terms of the wing surface downwash.
Revell, J. D.↗
Canard-wing lift interference related to maneuvering aircraft at subsonic speeds
An investigation was conducted at Mach numbers of 0.7 and 0.9 to determine the lift interference effect of canard location on wing planforms typical of maneuvering fighter configurations. The canard had an exposed area of 16.0 percent of the wing reference area and was located in the plane of the wing or in a position 18.5 percent of the wing mean geometric chord above the wing plane. In addition, the canard could be located at two longitudinal stations. Two different wing planforms were tested: one with a leading-edge sweep angle of 60 deg and the other with a leading-edge sweep angle of 44 deg. The results indicated that although downwash from the canard reduced the wing lift at angles of attack up to approximately 16 deg, the total lift was substantially greater with the canard on than with the canard off. At angles of attack above 16 deg, the canard delayed the wing stall. Changing canard deflection had essentially no effect on the total lift, since the additional lift generated by the canard deflection was lost on the wing due to an increased downwash at the wing from the canard.
Gloss, B. B.↗
Application of unsteady lifting surface theory to propellers in forward flight
Development of a theory for determining the aerodynamic forces for unsteady, compressible subsonic flow on a propeller in forward flight. The acceleration potential method is used in developing the basic downwash integral equation which governs the flow. This integral equation is solved by the doublet-lattice method, which consists of placing 'load' lines at certain locations on the chord and satisfying the downwash condition at other selected positions. The examples presented include the spanwise and chordwise loading on a rotating propeller for incompressible flow, an example of compressible flow calculations, and, finally, a calculation illustrating the loss of aerodynamic damping of a propeller blade due to the passage of the blade over its own wake.
Hammond, C. E.↗
Wind tunnel investigation of helicopter rotor wake effects on three helicopter fuselage models
The effects of rotor downwash on helicopter fuselage aerodynamic characteristics were investigated. A rotor model for generating the downwash was mounted close to each of three fuselage models. The main report presents the force and moment data in both graphical and tabular form and the pressure data in graphical form. This supplement presents the pressure data in tabular form. Each run or parameter sweep is identified by a unique run number. The data points in each run are identified by a point number. The pressure data can be matched to the force data by matching the run and point number.
Wilson, J. C.↗