A Method for Predicting the Elevator Deflection Required to Land
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EQUATIONS - EQUILIBRIUM↗
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Wind-tunnel vibration tests of a four-blade single rotating pusher propeller
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PROPELLERS - HAMILTON STANDARD CONSTANT-SPEED (HYD↗
Preliminary Results of an Investigation by the Wing-Flow Method of the Longitudinal Stability Characteristics of a 1/50-Scale Semispan Model of the McDonnell XP-88 Airplane
This paper presents the results of measurements of longitudinal stability of a 1/50-scale model of the XP-88 airplane by the wing-flow method. Lift, rolling-moment, hinge-moment, and pitching-moment characteristics as well as the downwash at the tail were measured over a Mach number range from approximately 0.5 to 1.05 at Reynolds numbers below 1,000,000. No measurements of drag were obtained. No abrupt changes due to Mach number were noted in any of the parameters measured. The data indicated that the wing was subject to early tip stalling; that the tail effectiveness decreased gradually with increasing Mach number up to M = 0.9, but increased again at higher Mach numbers; that the variation of downwash with angle of attack did not change appreciably with Mach number except between 0.95 and 1.0 where d(epsilon)/d(alpha), decreased from 0.46 to 0.32; that at zero lift with a stabilizer setting of -1.5 deg there was a gradually increasing nosing-up tendency with increasing Mach number; and that the control-fixed stability in maneuvers at constant speed gradually increased with increasing Mach number.
Crane, Harold L.↗
Space shuttle: S-IC booster/Grumman C4 orbiter determination of downwash on 900 square feet, 30 deg oriented S-IC fins and optimum orbiter body and aerodynamic surface incidence angles
Wind tunnel tests of downwash aerodynamic surface incidence angles on Saturn S-1C booster coupled to Grumman C-4 orbiter shuttle
Shrout, L. C.↗
Tables for the Rapid Estimation of Downwash and Sidewash Behind Wings Performing Various Motions at Supersonic Speeds
Equations for the downwash and sidewash due to supersonic yawed and unswept horseshoe vortices have been utilized in formulating tables and charts to permit a rapid estimation of the flow velocities behind wings performing various steady motions. Tabulations are presented of the downwash and sidewash in the wing vertical plane of symmetry due to a unit-strength yawed horseshoe vortex located at 20 equally spaced spanwise positions along lifting lines of various sweeps. (The bound portion of the yawed vortex is coincident with the lifting line.) Charts are presented for the purpose of estimating the spanwise variations of the flow-field velocities and give longitudinal variations of the downwash and sidewash at a nuMber of vertical and spanwise locations due to a unit-strength unswept horseshoe vortex. Use of the tables and charts to calculate wing downwash or sidewash requires a knowledge of the wing spanwise distribution of circulation. Sample computations for the rolling sidewash and angle-of-attack downwash behind a typical swept wing are presented to demonstrate the use of the tables and charts.
Bobbitt, Percy J.↗
Investigation of Downwash, Sidewash, and Mach Number Distribution behind a Rectangular Wing at a Mach Number of 2.41
An investigation of the nature of the flow field behind a rectangular circular-arc wing has been conducted in the Langley 9-inch supersonic tunnel. Pitot- and static-pressure surveys covering a region of flow behind the wing have been made together with detailed pitot surveys throughout the region of the wake. In addition, the flow direction has been measured using a weathercocking vane measurements. Theoretical calculations of the variation of both downwash and sidewash with angle of attack using Lagerstrom's superposition method have been made. In addition the effect of the wing thickness on the sidewash with the wing at 0 angle of attack has been evaluated. Near an angle of attack of 0, agreement between theory and experiment is good, particularly for the downwash results, except in the plane of the wing, inboard of the tip. In this region the proximity of the shed vortex sheet and the departure of the spanwise distribution of vorticity from theory would account for the disagreement. At higher angles of attack prediction of downwash depends on a knowledge of the location of the trailing vortex sheet, in order that the downwash may be corrected for its displacement and distortion. The theoretical location of the trailing vortex sheet, based on the theoretical downwash values integrated downstream from the wing trailing edge, is shown to differ widely from the experimental case. The rolling-up of the trailing vortex sheet behind the wing tip is evidenced by both the wake surveys and the flow-angle measurements.
Adamson, D.↗
Full-Scale Wind-Tunnel Tests of a PCA-2 Autogiro Rotor
This report presents the results of force tests on and air-flow surveys near PCA-2 autogiro rotor in the NACA full-scale wind tunnel. The force tests were made at three pitch settings and several rotor speeds; the effect of fairing protuberances on the rotor blade was determined. Induced downwash and yaw angles were determined at low tip-speed ratios in a plane 1 1/2 feet above the path of the blade tips. The results show that the maximum l/d of the rotor cannot be appreciably increased by increasing the blade pitch angle above about 4.5 degrees at the blade tip; that the protuberances on the blades cause more than 5 percent of the total rotor drag; and that the rotor center-of-pressure travel is very small.
Wheatley, John B↗
Measurements of Downwash and Sidewash Behind Cruciform Triangular Wings at Mach Number 1.4
The results of an experimental investigation of the downwash and sidewash flow angles in a transverse plane behind a cruciform-wing and body combination are presented. The wing-body combination was composed of a cruciform arrangement of two identical triangular wings of aspect ratio 2.31 corresponding to a leading-edge sweep angle of 60°, placed 'on a body of fineness ratio of 16. The investigation was conducted at a Mach number of 1.40 and at a Reynolds number of 1.25 million based on the mean aerodynamic chord of the wings. Dawnwash and sidewash measurements were made for angles of bank of O°, 11-1/4°, 22-1/2°, and 45° and through an angle-of-attack range of ±8O°.
Benton E. Wetzel↗
Investigation of Downwash, Sidewash, and Mach Number Distribution Behind a Rectangular Wing at a Mach Number of 2.41
An investigation of the nature of the flow field behind a rectangular wing of circular arc cross section has been conducted in the Langley 9-inch supersonic tunnel. Pitot- and static-pressure surveys covering a region of flow behind the wing have been made together with detailed pitot surveys throughout the region of the wake. In addition, the flow direction has been measured by means of a weathercocking vane. Theoretical calculations have been made to obtain the variation of both downwash and sidewash with angle of attack by using the superposition method of Lagerstrom, Graham, and Grosslight. In addition, the effect of wing thickness on the sidewash with the wing at 0 degree angle of attack has been evaluated.
Adamson, David↗
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.↗
Towards an advanced vortex flap system: The cavity flap
An extension of the vortex flap concept was explored with the aim of providing high-alpha flight control capability coupled with maneuver drag reduction for highly swept wing configurations. A retractable lower surface flap mounted on a translating hinge is proposed, allowing chordwise extension as well as deflection, the two movements being independently controlled. The frontal cavity formed by the partially extended and deflected flap captures a vortex above a certain angle of attack. The cavity vortex downwash alleviates the effective incidence of the wing leading edge, thus modulating vortex lift; at the same time, the induced suction in the cavity generates thrust. These postulated aerodynamic features of the cavity flap were validated through low speed tunnel pressure and visualization tests on a 65 deg swept oblique wing model, which also provided initial trends of the leading edge vortex alleviation and cavity suction with respect to flap extension, deflection and angle of attack. Force tests on a 60 deg delta model further showed the cavity flap L/D performance to compare favorably with the conventional vortex flap. A two segment flap arrangement with independently control led segments was envisaged for exploiting the vortex modulation capability of the cavity flap for pitch, roll and yaw control, in addition to drag reduction at high angles of attack.
Rao, D. M.↗
Spatial resolution and downwash velocity corrections for multiple-hole pressure probes in complex flows
Correction schemes for finite spatial resolution and induced downwash velocity are presented which have application to the measurement of complex three-dimensional flow fields using five-hole angle-type pressure probes. In the study, induced downwash velocity is assumed to be proportional to the transverse gradients of streamwise velocity. The present correction schemes are validated by application to flows including vortices embedded within turbulent boundary layers and flows in a curved channel with 1.27-cm width, a 40-to-1 aspect ratio, and 59.7 cm of convex surface curvature.
Ligrani, P. M.↗
Investigation of Some Wake Vortex Characteristics of an Inclined Ogive-Cylinder Body at Mach Number 2
For a body consisting of a fineness-ratio-3 ogival nose tangent to a cylindrical afterbody 7.3 diameters long, pitot-pressure distributions in the flow field, pressure distributions over the body, and downwash distributions along a line through the vortex centers have been measured for angles of attack to 20 degrees. The Reynolds numbers, based on body diameter, were 0.15 x 10 to the 6th power and 0.44 x 10 to the 6th power. Comparisons of computed and measured vortex paths and downwash distributions are made. (author)
Jorgensen, Leland H↗
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↗
Pressure distributions and oil-flow patterns for a swept circulation-control wing
Pressure distributions and photographs of oil flow patterns are presented for a circulation control wing. The model was an aspect ratio four semispan wing mounted on the side wall of the NASA Ames Transonic Wind Tunnel. The airfoil was a 20 percent thick ellipse, modified with circular leading and trailing edges of 4 percent radius, and had a 25.4 cm constant chord. This configuration does not represent a specific wing design, but is generic. A full span, tangetial, rearward blowing, circulation control slot was incorporated ahead of the trailing edge on the upper surface. The wing was tested at Mach numbers from 0.3 to 0.75 at sweep angle of 0 to 45 deg with internal to external pressure ratios of 1.0 to 3.0. Lift and pitching momemt coefficients were obtained from measured pressure distributions at five span stations. When the conventional corrections resulting from sweep angle are applied to the lift and moment of circulation control sections, no additional corrections are necessary to account for changes in blowing efficiency. This is demonstrated for an aft sweep angle of 45 deg. An empirical technique for estimating the downwash distribution of a swept wing was validated.
Keener, Earl R.↗
Computational Aerodynamic Analysis in Support of the CRM Tail Cone Thruster Configuration Wind Tunnel Test
NASA’s Advanced Air Transport Technology (AATT) project is breaching the boundaries of aircraft design in pursuit of eco-friendly solutions that are compatible with urban noise comfort levels. Boundary layer ingesting (BLI) propulsion systems promise to reduce fuel burn with additional potential benefits in noise reduction. Type-II BLI systems of the STARC-ABL type are the subject of a test campaign planned for fiscal year 2022 in the National Transonic Facility (NTF), for which a CRM-based model with a retrofitted tail cone thruster (TCT) has been designed. The present work is a precursor to the NTF test, where the 240 cases planned for the experiment were simulated using the Launch, Ascent and Vehicle Aerodynamics (LAVA)computational framework. Solution sensitivity to angle of attack, engine operating conditions, and the presence of the supporting structure (sting) in the wind tunnel test are analyzed in the extensive dataset. The main flow features contributing to the inlet distortion are identified as the vertical tail wake, wing downwash and fuselage upsweep vortices, with the latter two experiencing the greatest sensitivity to angle of attack. Finally, results from an inlet-guide-vane(IGV) design/integration study are presented. The LAVA team and the turbomachinery design team at NASA’s Ames and Glenn Research Centers are collaborating in an effort to reduce flow distortion upstream of the fan by means of integrating an IGV system into the CRM-TCT model. A significant improvement in flow distortion metrics has been achieved since the initial design iteration. Results employing an actuator zone model with realistic radially-varying thrust profiles to simulate first-order fan effects within LAVA are presented.
AATT↗
Navier-Stokes simulations of tip vortices for fixed and rotating helicopter blades
Flowfield and tip vortex results are presented for a hovering rotor blade at subcritical and supercritical flow conditions for both nonlifting and lifting configurations. These results are calculated numerically by solving, in a time-accurate fashion, the unsteady thin layer Navier-Stokes equations written in rotor coordinates. The lifting calculations use an induced downwash correction, estimated from a free wake analysis, to the geometric angle of attack of the blades to account for the wake effect. Comparison of numerical results with the experimental data shows very good agreement for all cases considered. Alternate methods of calculating hovering rotor flowfield as steady state flowfield on isolated fixed-blade that have the same circulation distribution as that of rotor in hover are also explored.
Srinivasan, G. R.↗
Ice Accretion Prediction for a Typical Commercial Transport Aircraft
Ice accretion calculations were made for a modern commercial transport using the NASA Lewis LEWICE3D ice accretion code. The ice accretion calculations were made for the wing and horizontal tail using both isolated flow models and flow models incorporating the entire airplane. The isolated flow model calculations were made to assess the validity of using these simplified models in lieu of the entire model in the ice accretion analysis of full aircraft. Ice shapes typifying a rime and a mixed ice shape were generated for a 30 minute hold condition. In general, the calculated ice shapes looked reasonable and appeared representative of a rime and a mixed ice conditions. The isolated flow model simplification was good for the main wing except at the root where it overpredicted the amount of accreted ice relative to the full aircraft flow model. For the horizontal tail the size and amount of predicted ice compared well for the two flow models, but the position of the accretions were more towards the upper surface for the aircraft flow model relative to the isolated flow model. This was attributed to downwash from the main wing which resulted in a lower effective angle of attack for the aircraft tail.
Bidwell, C. S.↗