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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

The Effect of Moment of Area Rule Modifications on the Drag, Lift and Pitching Moment Characteristics of an Unswept Aspect Ratio 6 Wing and Body Combination

An experimental investigation was conducted to determine the effect of moment-of-area-rule modifications on the drag, lift, and pitching-moment characteristics of a wing-body combination with a relatively high aspect-ratio unswept wing. The basic configuration consisted of an aspect-ratio-6 wing with a sharp leading edge and a thickness ratio of 0.06 mounted on a cut-off Sears-Haack body. The model with full moment-of-area-rule modifications had four contoured pods mounted on the wing and indentations in the body to improve the longitudinal distributions of area and moments of area. Also investigated were modifications employing pods and indentations that were only half the size of the full modifications and modifications with partial body indentations. The models were tested at angles of attack from -2 deg to +12 deg at Mach numbers from 0.6 to 1.4. In general, the moment-of-area-rule modifications had a large effect on the drag characteristics of the models but only a small effect on their lift and pitching-moment characteristics. The modifications provided substantial reductions in the zero-lift drag at transonic and low supersonic speeds, but at subsonic speeds the drag was increased. Near Mach number 1.0, the model with full modification provided the greatest reduction in drag, but at the highest test Mach numbers the half modification gave the largest drag reduction. In general, the percent reductions of zero- lift drag obtained with the aspect-ratio-6 wing were as great or greater than those previously obtained with aspect-ratio-3 wings. The effect of the modifications on the drag due to lift was small except at Mach num- bers below 0.9 where the modified models had higher drag-rise factors. Above Mach number 0.9, the modified models had higher lift-drag ratios than the basic model. The modified models also had higher lift curve slopes and generally were slightly more stable than the basic configuration.

Dickey, Robert R.↗
NASA NTRS1959

Effects of Forebody Deflection on the Stability and Control Characteristics of a Canard Airplane Configuration with a High Trapezoidal Wing at a Mach Number of 2.01

An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.01 to determine the effects of forebody deflection on the stability and control characteristics of a canard airplane configuration. The configuration had a high trapezoidal aspect-ratio-3 wing, a trapezoidal canard surface, and a single swept vertical tail. Forebody deflection angles of 0 deg, 2 deg and deg were investigated. The results indicated that nose-up deflections of the forebody provided positive increments of pitching moment with little increase in drag and hence would be useful in reducing the pitch-control requirements and the attendant losses in lift-drag ratio due to trimming. Deflection of the forebody, however, aggravated the decrease in directional stability with increasing angle of attack by causing a loss in tail contribution and by increasing the instability of the wing-body combination.

Spearman, M. Leroy↗
NASA NTRS1961

Static Longitudinal Characteristics at High Subsonic Speeds of a Complete Airplane Model with a Highly Tapered Wing having the 0.80 Chord Line Unswept and with Several Tail Configurations

An investigation was made at high subsonic speeds of a complete model having a highly tapered wing and several tail configurations. The basic aspect-ratio-4.00 wing had zero taper and an unswept 0.80 chord line. Several aspect-ratio modifications to the basic wing were made by clipping off portions of the wing tips. The complete model was tested with a chord-plane tail, a T-tail, and a biplane tail (combined T-tail and chord-plane tail). The model was tested in the Langley high-speed 7- by 10-foot tunnel at Mach numbers from 0.60 to 0.92. The data show that, when reduced to the same static margin, all the tail configurations tested on the model provided fairly good stability characteristics, the biplane tail giving the.best overall characteristics as regards pitching-moment linearity. Changes in static margin at zero lift coefficient with Mach number were small for the model with these tails over the Mach number range investigated.

Goodson, Kenneth W.↗
NASA NTRS1961

Static Lateral Characteristics at High Subsonic Speeds of a Complete Airplane Model with a Highly Tapered Wing having the 0.80 Chord Line Unswept and with Several Tail Configurations

An investigation was made at high subsonic speeds of a complete model having a highly tapered wing and several tail configurations. The aspect-ratio-3.50 wing had a taper ratio of 0.067 and an unswept 0.80 chord line. The complete model was tested with a wing-chord-plane tail, a T-tail, and a biplane tail (combined T-tail and wing-chord-plane tail). The model was tested in the Langley high-speed 7- by 10-foot tunnel at Mach numbers from 0.60 to 0.92 over a range of angle of attack of about +/- 20 deg. and a range of sideslip of -15 deg. to 13 deg. Some data were obtained with the horizontal stabilizer deflected. A few tests were also made with the wing tips clipped to an aspect ratio of 3.00. The data show that shock-interference effects between the tail surfaces (T-tail) can have considerable influence on the directional stability and effective dihedral. For example, the T-tail configuration with horizontal-tail leading-edge overhang showed a considerable loss in directional stability as the angle of attack was reduced to zero or negative values; whereas, the T-tail with zero leading-edge overhang showed the loss to be considerably less. The directional stability of the model with the low tail was essentially constant over a range of angle of attack of +/- 50 deg. All configurations tested showed a large reduction in stability at positive and negative angles of attack larger than about 15 deg., probably because of adverse sidewash associated with wing stall. The data show that a wing-chord-plane horizontal tail (low tail) tends to give a positive pitching-moment increment with increase in sideslip angle; whereas, a high tail (T-tail) tends to give negative increments in pitching moment.

Goodson, Kenneth W.↗
NASA NTRS1955

An Experimental Investigation of the Hinge-moment Characteristics of a Constant-chord Control Surface Oscillating at High Frequency

The results of an experimental investigation of the hinge-moment characteristics of a constant-chord control surface oscillating at high frequency is presented. The control surface was mounted on an aspect-ratio-2 triangular wing. The aerodynamic restoring-moment coefficient and damping-moment coefficient were determined at a frequency of 260 cycles per second for a Mach number range of 0.6 to 0.8 and 1.3 to 1.9 at angles of attack of 5 degrees and 10 degrees. The test results showed linear theory to be a reliable guide to the prediction of the trend of the restoring-moment coefficient with Mach number for the supersonic speed range of the investigation but overestimated the magnitude of the coefficient. The experimental values of the damping-moment coefficient were, for the most part, more positive than those indicated by the theory and, for some conditions, could lead to instability of the control surface. Comparison of the results of this investigation with those of previous investigations at 0 and 50 cycles per second showed that frequency had little effect on the restoring-moment coefficient. The damping-moment coefficient was similarly insensitive to frequency at an oscillation amplitude of plus-or-minus 1.0 degrees but at an amplitude ofplus-or-minus 2.5 degrees the results showed a destabilizing shift with increasing frequency.

Reese, David E JR↗
NASA NTRS1967

Subsonic and Supersonic Flutter Analysis of a Highly Tapered Swept-Wing Planform, Including Effects of Density Variation and Finite Wing Thickness, and Comparison with Experiments

The flutter characteristics of several wings with an aspect-ratio of 4.0, a taper ratio of 0.2, and a quarter-chord sweepback of 45 deg. have been investigated analytically for Mach numbers up to 2.0. The calculations were based on the modified-strip-analysis method, the subsonic-kernel-function method, piston theory, and quasi-steady second-order theory. Results of t h e analysis and comparisons with experiment indicated that: (1) Flutter speeds were accurately predicted by the modified strip analysis, although accuracy at t h e highest Mach numbers required the use of nonlinear aerodynamic theory (which accounts for effects of wing thickness) for the calculation of the aerodynamic parameters. (2) An abrupt increase of flutter-speed coefficient with increasing Mach number, observed experimentally in the transonic range, was also indicated by the modified strip analysis. (3) In the low supersonic range for some densities, a discontinuous variation of flutter frequency with Mach number was indicated by the modified strip analysis. An abrupt change of frequency appeared experimentally in the transonic range. (4) Differences in flutter-speed-coefficient levels obtained from tests at low supersonic Mach numbers in two wind tunnels were also predicted by the modified strip analysis and were shown to be caused primarily by differences in mass ratio. (5) Flutter speeds calculated by the subsonic-kernel-function method were in good agreement with experiment and with the results of the modified strip analysis. (6) Flutter speed obtained from piston theory and from quasi-steady second-order theory were higher than experimental values by at least 38 percent.

Yates, Carson, Jr.↗
NASA NTRS1957

Hydrodynamic characteristics over a range of speeds up to 80 feet per second of a rectangular modified flat plate having an aspect ratio of 0.25 and operating at several depths of submersion

Results of an investigation of the hydrodynamic characteristics over an extended speed range of a rectangular modified flat plate having an aspect ratio of 0.25 and operating at several depths of submersion are presented. Comparisons between these data and data over a lower speed range on a similar aspect-ratio-0.25 flat plate but having one-half the thickness are presented. These comparisons show no significant differences at the low speeds. At high speeds and high angles of attack, where extensive cavitation was present, the lift coefficients were lower than would have been indicated by the results of the previous investigations and the present investigation at the lower angles of attack. A brief discussion and comparison of ventilation are presented which shows two types of planing bubble formation and the effect of increasing the thickness of the model on the ventilation boundary.

REYNOLDS NUMBER EFFECTS - COMPLETE WINGS↗
NASA NTRS1958

Large-Scale Wind-Tunnel Tests of an Airplane Model with an Unswept, Aspect-Ratio-10 Wing, Two Propellers, and Blowing Flaps

An investigation was made to determine the lifting effectiveness and flow requirements of blowing over the trailing-edge flaps and ailerons on a large-scale model of a twin-engine, propeller-driven airplane having a high-aspect-ratio, thick, straight wing. With sufficient blowing jet momentum to prevent flow separation on the flap, the lift increment increased for flap deflections up to 80 deg (the maximum tested). This lift increment also increased with increasing propeller thrust coefficient. The blowing jet momentum coefficient required for attached flow on the flaps was not significantly affected by thrust coefficient, angle of attack, or blowing nozzle height.

Griffin, Roy N., Jr.↗
NASA NTRS1960

Low-Speed Investigation of a Full-Span Internal-Flow Jet-Augmented Flap on a High-Wing Model with a 35 deg Swept Wing of Aspect Ratio 7.0

An investigation of a full-span 17-percent-chord internal-flow jet-augmented flap on an aspect-ratio-7.0 wing with 35 deg of sweepback has been made in the Langley 300-MPH 7- by 10-foot tunnel. Blowing over the conventional elevator and blowing down from a nose jet were investigated as a means of trimming the large diving moments at the high momentum and high lift coefficients. The results of the investigation showed that the model with the horizontal tail 0.928 mean aerodynamic chord above the wing-chord plane was stable to the maximum lift coefficient. The large diving-moment coefficients could be trimmed either with a downward blowing nose jet or by blowing over the elevator. Neither the downward blowing nose jet nor blowing over the elevator greatly affected the static longitudinal stability of the model. Trimmed lift coefficients up to 8.8 with blowing over the elevator and up to 11.4 with blowing down at the nose were obtained when the flap was deflected 70 deg and the total momentum coefficients were 3.26 and 4.69.

Turner, Thomas R.↗
NASA NTRS1961

STOL Characteristics of a Propeller-Driven, Aspect-Ratio-10, Straight-Wing Airplane with Boundary-Layer Control Flaps, as Estimated from Large-Scale Wind-Tunnel Tests

A study is presented of the improvements in take-off and landing distances possible with a conventional propeller-driven transport-type airplane when the available lift is increased by propeller slipstream effects and by very effective trailing-edge flaps and ailerons. This study is based on wind-tunnel tests of a 45-foot span, powered model, with BLC on the trailing-edge flaps and controls. The data were applied to an assumed airplane with four propellers and a wing loading of 50 pounds per square foot. Also included is an examination of the stability and control problems that may result in the landing and take-off speed range of such a vehicle. The results indicated that the landing and take-off distances could be more than halved by the use of highly effective flaps in combination with large amounts of engine power to augment lift (STOL). At the lowest speeds considered (about 50 knots), adequate longitudinal stability was obtained but the lateral and directional stability were unsatisfactory. At these low speeds, the conventional aerodynamic control surfaces may not be able to cope with the forces and moments produced by symmetric, as well as asymmetric, engine operation. This problem was alleviated by BLC applied to the control surfaces.

Weiberg, James A↗
NASA NTRS1959

An Experimental Investigation of the Effect of a Canard Control on the Lift, Drag, and Pitching Moment of an Aspect-Ratio 2.0 Triangular Wing Incorporating a Form of Conical Camber

The results of an experimental investigation to determine the effect of a canard control on the lift, drag, and pitching-moment characteristics of an aspect-ratio-2.0 triangular wing incorporating a form of conical camber are presented. The canard had a triangular plan form of aspect ratio 2.0 and was mounted in the extended chord plane of the wing. The ratio of the area of the exposed canard panels to the total wing area was 6.9 percent, and the ratio of the total areas was 12.9 percent. Data were obtained at Mach numbers from 0.70 to 2.22 through an angle-of-attack range from -6 deg to +18 deg with the canard on, and with the canard off. To provide a basis for comparison, the canard was also tested with a symmetrical wing having the same plan form, aspect ratio, and thickness distribution as the cambered wing. The results of the investigation showed that at the high subsonic speeds the gain in maximum lift-drag ratio achieved by camber was considerably reduced by the addition of a canard. At the supersonic speeds, the addition of the canard did not change the effect of camber on the maximum lift-drag ratios.

Menees, Gene P.↗
NASA NTRS1958

Lift, Drag, and Pitching Moment of an Aspect-Ratio-2 Triangular Wing with Leading-Edge Flaps Designed to Simulate Conical Camber

An investigation was conducted to determine the effectiveness of leading-edge flaps in reducing the drag at lifting conditions of a triangular wing of aspect ratio 2.0. The flaps, deflected to simulate conically cambered wings having a wide range of design lift coefficients, were tested over a Mach number range of 0.70 to 2.22 through an angle-of-attack variation from -6 deg to +18 deg at a constant Reynolds number of 3.68 million based on the wing mean aerodynamic chord. A symmetrical wing of the same plan form and aspect ratio was also tested to provide a basis for comparison. The experimental results showed that with the flaps in the undeflected position, a small amount of fixed leading-edge droop incorporated over the outboard 5 percent of the wing semispan was as effective at high subsonic speeds as conical camber in improving the maximum lift-drag ratio above that of the symmetrical wing. At supersonic speeds, the penalty in minimum drag above that of the symmetrical wing was less than that incurred by conical camber. Deflecting the leading-edge flaps about the hinge line through 80 percent of the wing semispan resulted in further improvements of the drag characteristics at lift coefficients above 0.20 throughout the Mach number range investigated. The lift and pitching-moment characteristics were not significantly affected by the leading-edge flaps.

Menees, Gene P.↗