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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1940

Drag Determination of the Forward Component of a Tricycle Landing Gear

Wind-tunnel tests were performed to determine the drag of the front-wheel arrangements of several types of tricycle landing gear. One wheel was tested in arrangements to simulate both nonretracted and partially retracted types. The landing gears were tested in conjunction with a fuselage, and the effects of wheel extension and longitudinal location were determined. The drag changed very little with either longitudinal location or wheel extension for the landing gear with the lowest drag; a completely faired landing gear of the wheelspan, single-strut type. The drag of the trouser-type landing gear increased considerably, however, with an increase in the wheel extension. The wheel of the unaired retractable landing gear was at least one-half retracted into the fuselage before the drag became less than that of the best nonretracted landing gear. The drag per unit frontal area of the landing gears of the present tests was about the same as that found for similar landing gears in earlier tests.

Harmon, Hubert N↗
NASA NTRS1951

Effect of Tail Surfaces on the Base Drag of a Body of Revolution at Mach Numbers of 1.5 and 2.0

Wind-tunnel tests were performed at Mach numbers of 1.5 and 2.0 to investigate the influence of tail surfaces on the base drag of a body of revolution without boattailing and having a turbulent boundary layer. The tail surfaces were of rectangular plan form of aspect ratio 2.33 and has symmetrical, circular-arc airfoil section. The results of the investigation showed that the addition of these tail surfaces with the trailing edges at or near the body base incurred a large increase in the base-drag coefficient. For a cruciform tail having a 10-percent-thick airfoil section, this increase was about 70 percent at a Mach number of 1.5 and 35 percent at a Mach number of 2.0. As the trailing edge of the tail was moved forward or rearward of the base by about one tail-chord length, the base-drag increment was reduced to nearly zero. The increments in base-drag coefficient due to the presence of 10-percent-thick tail surfaces were generally twice those for 5-percent-thick surfaces. The base-drag increments due to the presence of a cruciform tail were less than twice those for a plane tail. An estimate of the change in base pressure due to the tail surfaces was made, based on a simple superposition of the airfoil-pressure field onto the base-pressure field behind the body. A comparison of the results with the experimental values indicated that in most cases the trend in the variation of the base-drag increment with changes in tail position could be predicted by this approximate method but that the quantitative agreement at most tail locations was poor.

BOUNDARY LAYER - BODIES OF REVOLUTION↗
NASA NTRS1947

Calculations of the Supersonic Wave Drag of Nonlifting Wings with Arbitrary Sweepback and Aspect Ratio: Wings Swept Behind the Mach Lines

On the basis of a recently developed theory for finite sweptback wings at supersonic speeds, calculations of the supersonic wave drag at zero lift were made for a series of wings having thin symmetrical biconvex sections with untapered plan forms and various angles of sweepback and aspect ratios. The results are presented in a unified form so that a single chart permits the direct determination of the wave drag for this family of airfoils for an extensive range of aspect ratio and sweepback angle for stream Mach numbers up to a value corresponding to that at which the Mach line coincides with the wing leading edge. The calculations showed that in general the wave-drag coefficient decreased with increasing sweepback. At Mach numbers for which the Mach lines are appreciably ahead of the wing leading edge, the 'wave-drag coefficient decreased to an important extent with increases in aspect ratio or slenderness ratio. At Mach numbers for which the Mach lines approach the wing leading edge (Mach numbers approaching a value equal to the secant of the angle of sweepback), the wave-drag coefficient decreased with reductions in aspect ratio or slenderness ratio. In order to check the results obtained by the theory, a comparison was made with the results of tests at the Langley Memorial Aeronautical Laboratory of sweptback wing attached to a freely falling body. The variation of the drag with Mach number and aspect ratio as given by the theory appeared to be in reasonable

Harmon, Sidney M↗
NASA NTRS1956

Flight Determination of Drag of Normal-Shock Nose Inlets with Various Cowling Profiles at Mach Numbers from 0.9 to 1.5

External-drag data are presented for normal-shock nose inlets with NACA 1-series, parabolic, and conic cowling profiles. The tests were made at an angle of attack of 0 degrees by using rocket-propelled models in free flight at Mach numbers from 0.9 to 1.5. The Reynolds number based on body maximum diameter varied from 2.5 x 10 sup 6 to 5.5 x 10 sup 6. At maximum flow rate, the inlet models had about the same external drag at a Mach number of approximately 1.1, but at higher Mach numbers the sharp-lip conic cowling had the least drag. Blunting or beveling the lip of the conic cowling while keeping the fineness ratio constant resulted in drag coefficients slightly higher than for the sharp-lip conic cowling at maximum flow rate. At a mass-flow ratio of about 0.8, the conic cowlings with sharp, blunt, or beveled lips and the parabolic cowling all gave about the same drag. The higher drag of the NACA 1-49-300 cowling, compared with the blunt-lip conic cowling, is associated with the greater fullness back of the inlet.

Sears, R I↗
NASA NTRS1944

Intercooler cooling-air weight flow and pressure drop for minimum drag loss

An analysis has been made of the drag losses in airplane flight of cross-flow plate and tubular intercoolers to determine the cooling-air weight flow and pressure drop that give a minimum drag loss for any given cooling effectiveness and, thus, a maximum power-plant net gain due to charge-air cooling. The drag losses considered in this analysis are those due to (1) the extra drag imposed on the airplane by the weight of the intercooler, its duct, and its supports and (2) the drag sustained by the cooling air in flowing through the intercooler and its duct. The investigation covers a range of conditions of altitude, airspeed, lift-drag ratio, supercharger-pressure ratio, and supercharger adiabatic efficiency. The optimum values of cooling air pressure drop and weight flow ratio are tabulated. Curves are presented to illustrate the results of the analysis.

Reuter, J George↗
NASA NTRS1953

Pressure drag of bodies at Mach numbers up to 2.0

The drag of bodies has now assumed greater importance because, as shown in NACA RM L53I15a, 1953 and NACA RM A53H18a, 1953, the transonic drag rise of an airplane can be the same as its equivalent body. Obviously, the airplane designer would like his airplane to have a low-drag equivalent body. This paper shows some of the factors which minimize the drag of bodies at transonic and supersonic speeds and shows some of the penalties caused by deviating from low-drag body shapes. Drag reductions can be obtained in two ways, first, through increasing the body fineness ratio, and second, through better shaping of the body profile at a given fineness ratio. The effects of fineness ratio are discussed first and then, more completely, detail-shape effects.

Nelson, Robert L↗
NASA NTRS1954

An Experimental Investigation of Reduction in Transonic Drag Rise at Zero Lift by the Addition of Volume to the Fuselage of a Wing-body-tail Configuration and a Comparison with Theory

An experimental investigation was made by the free-fall recoverable-model technique to assess at zero lift the possibilities of reducing the drag-rise coefficients of a wing-body-cruciform-tail combination by adding volume to the fuselage. The basic features of the test model were an unswept aspect-ratio-3.1 thin wing, a fineness-ratio-12.4 fuselage, and four 45 degrees sweptback tail surfaces. The tests covered a Mach number range of 0.84 to 1.15 with Reynolds numbers of 6.000.000 to 14,000,000, based on the wing mean aerodynamic chord. Considerable reduction in drag-rise coefficient was effected for several different modifications by the addition of properly distributed volume to the fuselage. In one instance, a reduction in drag coefficient was obtained by adding a volume which was almost four times the exposed wing volume. The computation method presented in NACA RM A53H17 generally predicted the supersonic drag-rise coefficients for each modification within 20 percent of the experimental values. As in the above-mentioned report, the predictions at a Mach number of one were not accurate. The changes in drag-rise coefficients resulting from the modifications were generally predicted with better accuracy than the values of drag-rise coefficients.

Holdaway, George H↗
NASA NTRS1960

Some Examples of the Applications of the Transonic and Supersonic Area Rules to the Prediction of Wave Drag

The experimental wave drags of bodies and wing-body combinations over a wide range of Mach numbers are compared with the computed drags utilizing a 24-term Fourier series application of the supersonic area rule and with the results of equivalent-body tests. The results indicate that the equivalent-body technique provides a good method for predicting the wave drag of certain wing-body combinations at and below a Mach number of 1. At Mach numbers greater than 1, the equivalent-body wave drags can be misleading. The wave drags computed using the supersonic area rule are shown to be in best agreement with the experimental results for configurations employing the thinnest wings. The wave drags for the bodies of revolution presented in this report are predicted to a greater degree of accuracy by using the frontal projections of oblique areas than by using normal areas. A rapid method of computing wing area distributions and area-distribution slopes is given in an appendix.

Nelson, Robert L.↗
NASA NTRS1959

Hydrodynamic Characteristics of Two Low-Drag Supercavitating Hydrofoils

An experimental investigation has been conducted in Langley tank no. 2 to determine the hydrodynamic characteristics of two low-drag supercavitating hydrofoils operating in a range of cavitation numbers from 0 to approximately 6. The hydrofoils had aspect ratios of 1 and 3, and the sections were derived by assuming five terms in the vorticity-distribution expansion of the equivalent airfoil. The aspect-ratio-1 hydrofoil was also tested at zero cavitation number with two sets of end plates having depths of 3/8 and 1/4 chords. Zero cavitation number was established by operating the hydrofoils near the water surface so that complete ventilation of the upper surfaces could be obtained. For those depths of submersion where complete ventilation was not obtained through vortex ventilation, two probes were used to introduce air to the upper surfaces of the hydrofoils and to induce complete ventilation. Data were obtained for a range of speeds from 20 to 80 fps, angles of attack from 2 to 20 deg, and ratios of depth of submersion to chord from 0 to 0.85. The experimental results obtained from the aspect-ratio-1 and aspect-ratio-3, five-term hydrofoils were compared with a three-dimensional zero-cavitation-number theory. The theoretical and experimental values of lift and center of pressure for the aspect-ratio-1 hydrofoil were in agreement, within engineering accuracy, for the range of lift coefficients investigated. The theoretical drag coefficients were lower, by a constant amount, than the experimental drag coefficients. The theoretical expressions derived for the lift, drag, and center of pressure of the aspect-ratio-3 hydrofoil were in agreement, within engineering accuracy, with the experimental values. The theoretical and experimental drag coefficients of the aspect-ratio-3 five-term hydrofoil were lower than the theoretical drag coefficients computed for a comparable Tulin-Burkart hydrofoil.

McGehee, John R.↗
NASA NTRS1959

A Numerical Method for Calculating the Wave Drag of a Configuration from the Second Derivative of the Area Distribution of a Series of Equivalent Bodies of Revolution

A method based on linearized and slender-body theories, which is easily adapted to electronic-machine computing equipment, is developed for calculating the zero-lift wave drag of single- and multiple-component configurations from a knowledge of the second derivative of the area distribution of a series of equivalent bodies of revolution. The accuracy and computational time required of the method to calculate zero-lift wave drag is evaluated relative to another numerical method which employs the Tchebichef form of harmonic analysis of the area distribution of a series of equivalent bodies of revolution. The results of the evaluation indicate that the total zero-lift wave drag of a multiple-component configuration can generally be calculated most accurately as the sum of the zero-lift wave drag of each component alone plus the zero-lift interference wave drag between all pairs of components. The accuracy and computational time required of both methods to calculate total zero-lift wave drag at supersonic Mach numbers is comparable for airplane-type configurations. For systems of bodies of revolution both methods yield similar results with comparable accuracy; however, the present method only requires up to 60 percent of the computing time required of the harmonic-analysis method for two bodies of revolution and less time for a larger number of bodies.

Levy, Lionel L., Jr.↗
NASA NTRS1959

Approach and Landing Investigation at Lift-Drag Ratios of 2 to 4 Utilizing a Straight-Wing Fighter Airplane

A series of landings was performed with a straight-wing airplane to evaluate the effect of low lift-drag ratios on approach and landing characteristics. Landings with a peak lift-drag ratio as low as 3 were performed by altering the airplane configuration (extending speed brakes, flaps, and gear and reducing throttle setting). As lift-drag ratio was reduced, it was necessary either to make the landing pattern tighter or to increase initial altitude, or both. At the lowest lift-drag ratio the pilots believed a 270 deg overhead pattern was advisable because of the greater ease afforded in visually positioning the airplane. The values of the pertinent flare parameters increased with the reduction of lift-drag ratio. These parameters included time required for final flare; speed change during final flare; and altitude, glide slope, indicated airspeed, and vertical velocity at initiation of final flare. The pilots believed that the tolerable limit was reached with this airplane in the present configuration, and that if, because of a further reduction in lift-drag ratio, more severe approaches than those experienced in this program were attempted, additional aids would be required to determine the flare-initiation point.

Matranga, Gene J.↗
NASA NTRS2002

Summary of Data from the First AIAA CFD Drag Prediction Workshop

The results from the first AIAA CFD Drag Prediction Workshop are summarized. The workshop was designed specifically to assess the state-of-the-art of computational fluid dynamics methods for force and moment prediction. An impartial forum was provided to evaluate the effectiveness of existing computer codes and modeling techniques, and to identify areas needing additional research and development. The subject of the study was the DLR-F4 wing-body configuration, which is representative of transport aircraft designed for transonic flight. Specific test cases were required so that valid comparisons could be made. Optional test cases included constant-C(sub L) drag-rise predictions typically used in airplane design by industry. Results are compared to experimental data from three wind tunnel tests. A total of 18 international participants using 14 different codes submitted data to the workshop. No particular grid type or turbulence model was more accurate, when compared to each other, or to wind tunnel data. Most of the results overpredicted C(sub Lo) and C(sub Do), but induced drag (dC(sub D)/dC(sub L)(exp 2)) agreed fairly well. Drag rise at high Mach number was underpredicted, however, especially at high C(sub L). On average, the drag data were fairly accurate, but the scatter was greater than desired. The results show that well-validated Reynolds-Averaged Navier-Stokes CFD methods are sufficiently accurate to make design decisions based on predicted drag.

Levy, David W.↗
NASA NTRS2003

Evaluation of a Drag-Free Control Concept for Missions in Low Earth Orbit

Atmospheric drag causes the greatest uncertainty in the equations of motion for spacecraft in Low Earth Orbit (LEO). If atmospheric drag eflects can be continuously and autonomously counteracted through the use of a drag-fee control system, drag may essentially be eliminated from the equations of motion for the spacecraft. The main perturbations on the spacecraft will then be those due to the gravitational field, which are much more easily predicted Through dynamical analysis and numerical simulation, this paper presents some potential costs and benefits associated with the fuel used during continuous drag compensation. In light of this cost-benefit analysis, simulation results are used to validate the concept of drag-free control for LEO spacecraft missions having certain characteristics.

Fleck, Melissa E.↗
NASA NTRS1960

Effects of Fineness Ratio and Reynolds Number on the Low-Speed Crosswind Drag Characteristics of Circular and Modified-Square Cylinders

A wind-tunnel investigation has been made on modified-square and circular cylinders to determine the effects of fineness ratio and Reynolds numbers on the crosswind drag characteristics. Fineness ratios from 2 to 14 were investigated over a Reynolds number range from approximately 300,000 to 1,650,000 which corresponded to Mach numbers from 0.057 to 0.377.The result of the investigation show that at supercraft Reynolds numbers the drag coefficient of the circular cylinder increases with increasing Reynolds number for all fineness ratios but at low fineness ratios this effect is considerably less than at higher fineness ratios. For circular cylinders in the high fineness-ratio range there is a reduction in drag as the fineness ratio is decreased except for Reynolds numbers of 900,000 and 1,000,000, whereas at low fineness ratios the opposite trend generally occurs. The addition of hemispherical ends to the circular cylinder gave a substantial decrease in drag at a fineness ratio of 3.27 but the effect was negligible at fineness ratios of 5.27 and 10. The finite-length modified-square cylinder gave the reduction in drag over the two-dimensional modified-square cylinder for the complete range of test Reynolds numbers with the lowest fineness ratio giving the lowest drag at Reynolds numbers above 3O0,OOO.

McKinney, Linwood W.↗
NASA NTRS1999

Experimental Investigation of Convoluted Contouring for Aircraft Afterbody Drag Reduction

An experimental investigation was performed in the NASA Langley 16-Foot Transonic Tunnel to determine the aerodynamic effects of external convolutions, placed on the boattail of a nonaxisymmetric nozzle for drag reduction. Boattail angles of 15 and 22 were tested with convolutions placed at a forward location upstream of the boattail curvature, at a mid location along the curvature and at a full location that spanned the entire boattail flap. Each of the baseline nozzle afterbodies (no convolutions) had a parabolic, converging contour with a parabolically decreasing corner radius. Data were obtained at several Mach numbers from static conditions to 1.2 for a range of nozzle pressure ratios and angles of attack. An oil paint flow visualization technique was used to qualitatively assess the effect of the convolutions. Results indicate that afterbody drag reduction by convoluted contouring is convolution location, Mach number, boattail angle, and NPR dependent. The forward convolution location was the most effective contouring geometry for drag reduction on the 22 afterbody, but was only effective for M < 0.95. At M = 0.8, drag was reduced 20 and 36 percent at NPRs of 5.4 and 7, respectively, but drag was increased 10 percent for M = 0.95 at NPR = 7. Convoluted contouring along the 15 boattail angle afterbody was not effective at reducing drag because the flow was minimally separated from the baseline afterbody, unlike the massive separation along the 22 boattail angle baseline afterbody.

Deere, Karen A.↗
NASA NTRS1998

An Aerodynamic Assessment of Micro-Drag Generators (MDGs)

Commercial transports as well as fighter aircraft of the future are being designed with very low drag (friction and pressure). Concurrently, commuter airports are being built or envisioned to be built in the centers of metropolitan areas where shorter runways and/or reduced noise footprints on takeoff and landing are required. These requirements and the fact that drag is lower on new vehicles than on older aircraft have resulted in vehicles that require a large amount of braking force (from landing-gear brakes, spoilers, high-lift flaps, thrust reversers, etc.). Micro-drag generators (MDGs) were envisioned to create a uniformly distributed drag force along a vehicle by forcing the flow to separate on the aft-facing surface of a series of deployable devices, thus, generating drag. The devices are intended to work at any speed and for any type of vehicle (aircraft, ground vehicles, sea-faring vehicles). MDGs were applied to a general aviation wing and a representative fuselage shape and tested in two subsonic wind tunnels. The results showed increases in drag of 2 to 6 times that of a "clean" configuration.

Bauer, Steven X. S.↗
NASA NTRS1954

Effects of Fuselage Modifications on the Drag Characteristics of a 1/20-Scale Model of the Convair F-102 Airplane at Transonic Speeds

An investigation has been conducted in the Langley 8-foot transonic tunnel to determine the effects of several fuselage modifications on the transonic drag-rise characteristics of a 1/20-scale model of the Convair F-102 airplane. Tests covered an angle-of-attack range from 0deg to about 10deg and a Mach number range from 0.60 to 1.14. Results indicated that the transonic drag rise .for the basic F-102 airplane could be substantially reduced by extending the fuselage after-body approximately 8 percent of the fuselage length. Tests of other bodies indicated that a shorter (4-percent) afterbody extension may have a similar effect on the drag rise. Further improvement of the axial cross-sectional-area distribution of the 8-percent extended configuration through the addition of fuselage volume resulted in additional reductions in the drag rise at a Mach number of 1.0 and caused no or only slight drag penalties at the higher Mach numbers. The results of the present tests generally substantiate the area-rule concept with respect to the prediction of the transonic drag rise through the use of an equivalent-area body of revolution for a practical delta-wing airplane configuration.

Kelly, Thomas C.↗
NASA NTRS1955

Low-Lift Drag of the Grumman F9F-9 Airplane as Obtained by a 1/7.5-Scale Rocket-Boosted Model and by Three 1/45.85-Scale Equivalent-Body Models between Mach Numbers of 0.8 and 1.3, TED No. NACA DE 391

Low-lift drag data are presented herein for one 1/7.5-scale rocket-boosted model and three 1/45.85-scale equivalent-body models of the Grumman F9F-9 airplane, The data were obtained over a Reynolds number range of about 5 x 10(exp 6) to 10 x 10(exp 6) based on wing mean aerodynamic chord for the rocket model and total body length for the equivalent-body models. The rocket-boosted model showed a drag rise of about 0,037 (based on included wing area) between the subsonic level and the peak supersonic drag coefficient at the maximum Mach number of this test. The base drag coefficient measured on this model varied from a value of -0,0015 in the subsonic range to a maximum of about 0.0020 at a Mach number of 1.28, Drag coefficients for the equivalent-body models varied from about 0.125 (based on body maximum area) in the subsonic range to about 0.300 at a Mach number of 1.25. Increasing the total fineness ratio by a small amount raised the drag-rise Mach number slightly.

Stevens, Joseph E.↗