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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Unsteady surface pressure measurements on a slender delta wing undergoing limit cycle wing rock

An experimental investigation of slender wing limit cycle motion known as wing rock was investigated using two unique experimental systems. Dynamic roll moment measurements and visualization data on the leading edge vortices were obtained using a free to roll apparatus that incorporates an airbearing spindle. In addition, both static and unsteady surface pressure data was measured on the top and bottom surfaces of the model. To obtain the unsteady surface pressure data a new computer controller drive system was developed to accurately reproduce the free to roll time history motions. The data from these experiments include, roll angle time histories, vortex trajectory data on the position of the vortices relative to the model's surface, and surface pressure measurements as a function of roll angle when the model is stationary or undergoing a wing rock motion. The roll time history data was numerically differentiated to determine the dynamic roll moment coefficient. An analysis of these data revealed that the primary mechanism for the limit cycle behavior was a time lag in the position of the vortices normal to the wing surface.

Arena, Andrew S., Jr.↗
NASA NTRS1985

Bridle Attachment for Aircraft Spin-Recovery Parachute

Antispin rolling moment produced by chute drag force. Parachute stowed prior to deployment. At deployment, bridle attachment produces antispin rolling moment. At recovery, parachute forces are in aircraft plane of symmetry. Attachment system reduces parachute diameter typically required for spin recovery of experimental aircraft during harzardous flight testing.

White, W. L.↗
NASA NTRS1960

Supersonic Aerodynamic Characteristics of a Low-Drag Aircraft Configuration having an Arrow Wing of Aspect Ratio 1.86 and a Body of Fineness Ratio 20

A free-flight rocket-propelled-model investigation was conducted at Mach numbers of 1.2 to 1.9 to determine the longitudinal and lateral aero-dynamic characteristics of a low-drag aircraft configuration. The model consisted of an aspect-ratio -1.86 arrow wing with 67.5 deg. leading-edge sweep and NACA 65A004 airfoil section and a triangular vertical tail with 60 deg. sweep and NACA 65A003 section in combination with a body of fineness ratio 20. Aerodynamic data in pitch, yaw, and roll were obtained from transient motions induced by small pulse rockets firing at intervals in the pitch and yaw directions. From the results of this brief aerodynamic investigation, it is observed that very slender body shapes can provide increased volumetric capacity with little or no increase in zero-lift drag and that body fineness ratios of the order of 20 should be considered in the design of long-range supersonic aircraft. The zero-lift drag and the drag-due-to-lift parameter of the test configuration varied linearly with Mach number. The maximum lift-drag ratio was 7.0 at a Mach number of 1.25 and decreased slightly to a value of 6.6 at a Mach number of 1.81. The optimum lift coefficient, normal-force-curve slope, lateral-force-curve slope, static stability in pitch and yaw, time to damp to one-half amplitude in pitch and yaw, the sum of the rotary damping derivatives in pitch and also in yaw, and the static rolling derivatives all decreased with an increase in Mach number. Values of certain rolling derivatives were obtained by application of the least-squares method to the differential equation of rolling motion. A comparison of the experimental and calculated total rolling-moment-coefficient variation during transient oscillations of the model indicated good agreement when the damping-in-roll contribution was included with the static rolling-moment terms.

Gillespie, Warren, Jr.↗
NASA NTRS1975

Exploratory flight investigation of aircraft response to the wing vortex wake generated by the augmentor wing jet STOL research aircraft

A brief exploratory flight program was conducted at Ames Research Center to investigate the vortex wake hazard of a powered-lift STOL aircraft. The study was made by flying an instrumented Cessna 210 aircraft into the wake of the augmentor wing jet STOL research aircraft at separation distances from 1 to 4 n.mi. Characteristics of the wake were evaluated in terms of the magnitude of the upset of the probing aircraft. Results indicated that within 1 n.mi. separation the wake could cause rolling moments in excess of roll control power and yawing moments equivalent to rudder control power of the probe aircraft. Subjective evaluations by the pilots of the Cessna 210 aircraft, supported by response measurements, indicated that the upset caused by the wake of the STOL aircraft was comparable to that of a DC-9 in the landing configuration.

Jacobsen, R. A.↗
NASA NTRS1975

Theoretical study of lift-generated vortex wakes designed to avoid rollup

Two hypothetical vortex wakes are introduced and studied theoretically to explore whether the rollup of lift-generated vortex sheets can be suppressed. The circulation distribution across each wake is specified such that one rotates and the other translates as a unit due to their self-induced velocities. Several span loadings are constructed from these solutions and the resulting inviscid wake structure is computed for several span lengths behind the generating wing by use of the discrete vortex method wherein the vortex wake is represented by an array of vortices. The final distribution of vortices is then used to estimate the rolling moment on an encountering wing. It is found that, even though the initial specified motions are not sustained, substantial reductions in rolling moment are predicted for certain ranges of the ratio of the span of the generating wing to the following wing.

Rossow, V. J.↗
NASA NTRS1949

Aerodynamic Characteristics of a 0.5-Scale Model of the Fairchild XSAM-N-2 Lark Missile at High Subsonic Speeds

An investigation was conducted to determine the longitudinal- and lateral-stability characteristics of a 0.5-scale moue1 of the Fairchild Lark missile, The model was tested with 0 deg and with 22.5 deg of roll. Three horizontal wings having NACA 16-009, 16-209, and 64A-209 sections were tested. Pressures were measured on both pointed and blunt noses. The wind-tunnel-test data indicate that rolling the missile 22.5 deg. had no serious effect on the static longitudinal stability. The desired maneuvering acceleration could not be attained with any of the horizontal wings tested, even with the horizontal wing flaps deflected 50 deg. The flaps on the 64A-209 wing (with small trailing-edge angles and flat sides) were effective at all flap deflections, while the flaps on the 16-series wings (with large trailing-edge angles) lost effectiveness at small flap deflections. The data showed that rolling moment existed when the vertical wing flaps were deflected with the model at other than zero angle of attack. A similar rolling moment probably would be found . with the horizontal wing flaps deflected and the model yawed.

Martin, Andrew↗
NASA NTRS1977

The development and use of spoilers as vortex attenuators

Results are presented of groundbased and flight investigations performed to develop spoilers as trailing vortex alleviation devices. Based on the results obtained in these investigations, it was found that the induced rolling moment on a trailing model can be reduced by spoilers located near the midsemispan of a vortex generating wing. Substantial reductions in induced rolling moment occur when the spoiler vortex attenuator is located well forward on both unswept and swept wing models. In addition, it was found that existing flight spoilers on the jumbo-jet transport aircraft can be effective as trailing vortex attenuators.

Croom, D. R.↗
NASA NTRS1977

Wind-tunnel and flight evaluation of spoilers as trailing-vortex hazard alleviation devices

This paper presents the results of ground-based and flight investigations that have been performed at NASA for the purpose of development of spoilers as trailing-vortex hazard alleviation devices. Based on the results obtained in these investigations, it was found that the induced rolling moment on a trailing model can be reduced by spoilers located near the mid-semispan of a vortex-generating wing. Substantial reductions in induced rolling moment occur when the spoiler vortex attenuator is located well forward on both unswept and swept wing models. In addition, it was found by ground-based model tests and verified by full-scale flight tests that the existing flight spoilers on the B-747 aircraft are effective as trailing vortex attenuators. Based on the results of wind-tunnel investigations of the DC-10-30 and L-1011 aircraft models, the existing flight spoilers on both the DC-10-30 and L-1011 aircraft may also be effective trailing vortex attenuators.

Croom, D. R.↗
NASA NTRS1948

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↗
NASA NTRS1983

Aileron effectiveness for a subsonic transport model with a high-aspect-ratio supercritical wing

Aileron effectiveness for a subsonc energy efficient transport (EET) model with a high aspect ratio supercritical wing was determined in the 8-foot transonic pressure tunnel. Data are presented for ailerons located at three positions along the wing span. The ailerons were designed as a preliminary active control concept with gust load alleviation, maneuver load alleviation, and flutter suppression systems. A linear variation of rolling moment coefficient with angle of attack for individual and multiple aileron deflections at Mach numbers up to 0.81 is indicated. For Mach numbers greater than 0.81, the rolling moment coefficient data become nonlinear with increasing angle of attack. At Mach numbers near the design value increased aileron effectiveness resulted from aft transition locations, which produced relatively thin boundary layers and greater effective aileron deflections. Individual aileron deflections on the right wing panel produced only small effects on yawing moment and side force coefficients.

Jacobs, P. F.↗
NASA NTRS1986

Mutual downwash interference effects on the normal force and roll driving coefficients of multi-stage rockets

For several years, the TAD2 computer program has been used to calculate the aerodynamic properties of sounding rockets. This program was limited, however, in that a maximum of two fin sets could be accommodated analytically. In addition, the roll moment coefficients did not consider downwash effects. Techniques are developed and equations derived for the expansion of capability to include three fin sets. The effects of downwash on the roll moment coefficients are also treated analytically for two and three stage rockets.

Ward, P. R.↗
NASA NTRS1988

Research on aircraft/vortex-wake interactions to determine acceptable level of wake intensity

An evaluation of the literature on large aircraft wake-vortex encounters in flight and in flight simulators has furnished an estimate of the level to which the vortex-induced rolling moments must be reduced in order to be perceived as nonhazardous at a 2-n.mi. separation distance. The criteria are based on the ratio of the vortex-induced acceleration in roll to the aileron-induced roll acceleration. A wake is acceptably alleviated if the ratio of vortex-to-aileron rolling moments is less than about 0.5. When a satisfactory alleviation scheme is identified, the alleviated vortex structure should be inserted into a simulator to ascertain whether the maximum bank angles induced are within tolerable limits.

Rossow, Vernon J.↗
NASA NTRS1975

Experimental study of the effect on span loading on aircraft wakes

Measurements were made in the NASA-Ames 40- by 80-foot wind tunnel of the rolling moment induced on a following model in the wake 13.6 spans behind a subsonic transport model for a variety of trailing edge flap settings of the generator. It was found that the rolling moment on the following model was reduced substantially, compared to the conventional landing configuration, by reshaping the span loading on the generating model to approximate a span loading, found in earlier studies, which resulted in reduced wake velocities. This was accomplished by retracting the outboard trailing edge flaps. It was concluded, based on flow visualization conducted in the wind tunnel as well as in a water tow facility, that this flap arrangement redistributes the vorticity shed by the wing along the span to form three vortex pairs that interact to disperse the wake.

Corsiglia, V. R.↗
NASA NTRS1975

Experimental study of the effect of span loading on aircraft wakes

Measurements were made in the NASA-Ames 40- by 80-Foot Wind Tunnel of the rolling moment induced on a following model in the wake 13.6 spans behind a subsonic transport model for a variety of trailing edge flap settings of the generator. It was found that the rolling moment on the following model was reduced substantially, compared to the conventional landing configuration, by reshaping the span loading on the generating model to approximate a span loading, found in earlier studies, which resulted in reduced wake velocities. This was accomplished by retracting the outboard trailing edge flaps. It was concluded, based on flow visualization conducted in the wind tunnel as well as in a water tow facility, that this flap arrangement redistributes the vorticity shed by the wing along the span to form three vortex pairs that interact to disperse the wake.

Corsiglia, V. R.↗
NASA NTRS1976

Wind-Tunnel Investigation of the Effect of Porous Spoilers on the Wake of a Subsonic Transport Model

Tests were conducted in the Ames Research Center 40- by 80-Foot Wind Tunnel to determine how porosity of wing spoilers on a B-747 airplane would affect the rolling moments imposed on an aircraft following in the wake. It was found that spoilers with 40 percent porosity and hole diameter to thickness ratio of 1.1 were just as effective in reducing the rolling moment imposed on the follower as solid spoilers, for the case of two spoilers per wing panel (6.4 percent semispan each) with a following model whose span was 20 percent of the span of the generator. When a larger following model was tested, whose span was 50 percent of that of the generator, the effectiveness of the two spoilers per wing was substantially reduced.

Corsiglia, V. R.↗
NASA NTRS1954

Effect of Horizontal-Tail Span and Vertical Location on the Aerodynamic Characteristics of an Unswept Tail Assembly in Sideslip

An investigation has been conducted in the Langley stability tunnel on a vertical-tail model with a stub fuselage in combination with various horizontal tails to determine the effect of horizontal-tail span and vertical location of the horizontal tail relative to the vertical tail on the aerodynamic characteristics of an unswept tail assembly in sideslip. The results of the investigation indicated that the induced loading carried by the horizontal tail produced a rolling moment about the point of attachment to the vertical tail which was strongly influenced by horizontal-tail span and vertical locations. The greatest effect of horizontal-tail span on the rolling-moment derivative of the complete tail assembly was obtained for horizontal-tail locations near the top of the vertical tail. Span loadings which were reduced to the static-stability derivatives were calculated for each configuration tested by applying the well-known finite-step method used for wings to the intersecting surfaces of the vertical and horizontal tails. The finite-step method provides a simple and effective means of investigating the span loadings of intersecting surfaces.

Riley, Donald R↗
NASA NTRS1994

Experimental study at low supersonic speeds of a missile concept having opposing wraparound tails

A wind-tunnel investigation has been performed at low supersonic speeds (at Mach numbers of 1.60, and 2.16) to evaluate the aerodynamic characteristics of a missile concept capable of being tube launched and controlled with a simple one-axis canard controller. This concept, which features an axisymmetric body with two planar canards and four wraparound tail fins arranged in opposing pairs, must be in rolling motion to be controllable in any radial plane with the planar canards. Thus, producing a constant rolling moment that is invariant with speed and attitude to provide the motion is desirable. Two tail-fin shaping designs, one shaved and one beveled, were evaluated for their efficiency in producing the needed rolling moments, and the results showed that the shaved fins were much more desirable for this task than the beveled fins.

Allen, Jerry M.↗
NASA NTRS1978

Experimental investigation of wing fin configurations for alleviation of vortex wakes of aircraft

A variety of fin configurations were tested on a model of the Boeing B747 in 40 by 80 foot wind tunnels. The test results confirmed that a reduction in wake rolling moment was brought about by the vortex shed by the fins so that a wide range of designs can be used to achieve wake alleviation. It was also found that the reduction in wake-induced rolling moments was especially sensitive to the location of the smaller fins on the wing and that the penalties in lift and drag can probably be made negligible by proper fin design.

Rossow, V. J.↗