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At least 631 records · Page 35
All sourcesNASA NTRSDOE OSTI
NASA NTRS1960

Large-Scale Wind-Tunnel Tests of a Wingless Vertical Take-Off and Landing Aircraft: Preliminary Results

Large-scale wind-tunnel tests were made of a wingless vertical take-off and landing aircraft at zero sideslip to determine performance and longitudinal stability and control characteristics at airspeeds from 0 to 70 knots. Roll control and rudder effectiveness were also obtained. Limitations in the propulsion system restricted the lift for which level flight could be simulated to approximately 1500 pounds. Test variables with roll control and rudder undeflected were airspeed, vane setting, angle of attack, elevator deflection, and power. In most of the tests angle of attack, elevator, and power were varied individually while the other four parameters were held constant at previously determined values required for simulating trimmed level flight. The majority of the tests were made with power on and tail on at airspeeds between 20 and 70 knots. However, a limited number of data were obtained for the following conditions: (1) at zero velocity, horizontal tail on, power on; (2) at forward velocity, tail off and power on; and (3) at forward velocity, tail on, but with power off.

Koenig, David G.↗
NASA NTRS1960

Longitudinal Aerodynamic Characteristics of a Four-Propeller Deflected Slipstream VTOL Model Including the Effects of Ground Proximity

Results are presented of a wind-tunnel investigation of the longitudinal stability, control, and performance characteristics of a model of a four-propeller deflected-slipstream VTOL airplane in the transition speed range. These results indicate that steady level-flight transition and descending flight-path angles up to 7 or 8 deg. out of the region of ground effect can be accomplished without wing stall being encountered. In general, the pitching moments out of ground proximity can be adequately trimmed by programming the stabilizer incidence to increase with increasing flap deflection, except for a relatively large diving moment in the hovering condition. The deflection of the slipstream onto the horizontal tail in proximity of the ground substantially increases the diving moment in hovering, unless the tail is set at a large nosedown incidence.

Kuhn, Richard E.↗
NASA NTRS1960

Aerodynamic Characteristics at Mach Numbers of 1.41 and 2.01 of a Series of Cranked Wings Ranging in Aspect Ratio from 4.00 to 1.74 in Combination with a Body

A program has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel to determine the effects of certain wing plan-form variations on the aerodynamic characteristics of wing-body combinations at supersonic speeds. The present report deals with the results of tests of a family of cranked wing plan forms in combination with an ogive-cylinder body of revolution. Tests were made at Mach numbers of 1.41 and 2.01 at corresponding values of Reynolds number per foot of 3.0 x 10(exp 6) and 2.5 x 10(exp 6). Results of the tests indicate that the best overall characteristics were obtained with the low-aspect-ratio wings. Plan-form changes which involved decreasing the aspect ratio resulted in higher values of maximum lift-drag ratio, in addition to large increases in wing volume. Indications are that this trend would have continued to exist at aspect ratios even lower than the lowest considered in the present tests. Increases in the maximum lift-drag ratio of about 15 percent over the basic wing were achieved with practically no increase in drag. The severe longitudinal stability associated with the basic cranked wing was no longer present (within the limits of the present tests) on the wings of lower aspect ratio formed by sweeping forward the inboard portion of the trailing edge.

Sevier, John R., Jr.↗
NASA NTRS1960

The Longitudinal Aerodynamic Characteristics of a Sweptback Wing-Body Combination With and Without End Plates at Mach Numbers from 0.40 to 0.93

An investigation was made at high subsonic speeds in the Langley high-speed 7- by 10-foot tunnel to determine the effect of end plates on the longitudinal aerodynamic characteristics of a sweptback wing-body combination with and without drooped chord-extensions. The wing had 45 deg sweepback of the quarter-chord line, an aspect ratio of 4, a taper ratio of 0.3, and NACA 65AO06 airfoil sections parallel to the plane of symmetry, and was mounted near the rear of a body of revolution having a fineness ratio of approximately 8. The results indicated that the addition of the end plates to either the wing with drooped chord-extensions or to the wing without drooped chord-extensions slightly increased the lift in the low angle-of-attack range but slightly decreased the lift at moderate and high angles of attack. The addition of the end plates to the wing without the chord-extensions caused a small increase in the maximum lift-drag ratio at Mach numbers below 0.65 and a slight decrease at the higher Mach numbers; however, for the addition of the end plates to the wing with the chord- extensions the maximum lift-drag ratio was slightly decreased below a Mach number of 0.88, while a slight increase occurred for the higher Mach numbers. The addition of the end plates to the wings with and without the chord-extensions caused the static longitudinal stability to increase considerably for all Mach numbers; however, only a slight reduction in the aerodynamic-center variation with Mach number was observed.

Henderson, William P.↗
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 NTRS1959

Flight Studies of Problems Pertinent to High-Speed Operation of Jet Transports

A flight investigation was made to assess the potential operational problems of jet transports in the transonic cruise range. In this study a large multiengine jet airplane having geometric characteristics fairly representative of the jet transport was used; however, in order to ensure general applicability of the results, the aerodynamic characteristics of the test airplane were varied to simulate a variety of jet- transport airplanes. Some of the specific areas investigated include: (1) an overall evaluation of longitudinal stability and control characteristics at transonic speeds, with an assessment of pitch-up characteristics, (2) the effect of buffeting on airplane operational speeds and maneuvering, (3) the desirable lateral-directional damping characteristics, (4) the desirable lateral-control characteristics, (5) an assessment of over-speed and speed-spread requirements, including the upset maneuver, and (6) an assessment of techniques and airplane characteristics for rapid descent and slow-down. The results presented include pilots' evaluation of the various problem areas and specific recommendations for possible improvement of jet-transport operations in the cruising speed range.

Butchart, Stanley P.↗
NASA NTRS1958

Effect of Nose Length, Fuselage Length, and Nose Fineness Ratio on the Longitudinal Aerodynamic Characteristics of Two Complete Models at High Subsonic Speeds

An investigation has been made of the effects of nose length, fuselage length, and nose fineness ratio on the static longitudinal aerodynamic characteristics of an airplane model with a swept wing and low tail and of a second model with a highly tapered wing of moderate sweep and a T-tail. The tests were conducted in the Langley high-speed 7- by 10-foot tunnel at Mach numbers from 0.60 to 0.92. The nose and body cross sections were circular. For either the model with the swept wing and low tail or the model with the highly tapered wing of moderate sweep and the T-tail, the effects of forebody changes amounted primarily to rotations of the pitching-moment curves (changes in static margin) over the test ranges of angle of attack and Mach number. For the range of body shapes investigated the longitudinal stability at low lift is decreased by an increase in nose length or in fuselage length or by a reduction in nose fineness ratio when the fuselage length is held constant. In general, the stability for all model configurations showed substantially the same variation with changes in forebody area moment. The forebody changes did not alter the angle of attack at which an unstable break occurred in the moment contribution of the T-tail but did alter somewhat the magnitude of the instability.

Goodson, Kenneth W.↗
NASA NTRS1961

Aerodynamic Characteristics of Low-Aspect-Ratio Wings in Close Proximity to the Ground

A wind-tunnel investigation has been conducted to determine the effect of ground proximity on the aerodynamic characteristics of thick highly cambered rectangular wings with aspect ratios of 1. 2, 4, and 6. The results showed that, for these aspect ratios, as the ground war, approached all wings experienced increases in lift-curve slope and reductions in induced drag which resulted in increases in lift-drag ratio. Although an increase in lift-curve slope was obtained for all aspect ratios as the ground was approached, the lift coefficient at an angle of attack of 0 deg for any given aspect ratio remained nearly constant. The experimental results were in general agreement with Wieselsberger's ground-effect theory (NACA Technical Memorandum 77). As the wings approached the ground, there was an increase in static longitudinal stability at positive angles of attack. When operating in ground effect, all the wings had stability of height at positive angles of attack and instability of height at negative angles of attack. Wing-tip fairings on the wings with aspect ratios of 1 and 2 produced small increases in lift-drag ratio in ground effect. End plates extending only below the chord plane on the wing with an aspect ratio of 1 provided increases in lift coefficient and in lift-drag ratio in ground effect.

Fink, Marvin P.↗
NASA NTRS1958

Low-Speed Aerodynamic and Hydrodynamic Characteristics of a Proposed Supersonic Multijet Water-Based Hydro-Ski Aircraft with Upward-Rotating Engines

The low-speed aerodynamic and hydrodynamic characteristics of a proposed multijet water-based aircraft configuration for supersonic operation have been investigated. The design features include upward-rotating engines, body indentation, a single hydro-ski, and a wing with an aspect ratio of 3.0, a taper ratio of 0.143, 36.90 sweepback of the quarter-chord line, and NACA 65AO04 airfoil sections. For the aerodynamic investigation, with the flaps retracted, the model was longitudinally and directionally stable up to the stall. The all-movable horizontal tail was capable of trimming the model up to a lift coefficient of approximately 0.87. All flap configurations investigated had a tendency to become longitudinally unstable at stall. The effectiveness of the all-movable horizontal tail increased with increasing lift coefficient for all flap configurations investigated; however, with the large static margin of the configuration with the center of gravity at 0.25 mean aerodynamic chord, the all-movable horizontal tail was not powerful enough to trim all the various flapped configurations investigated throughout the angle-of-attack range. For the hydrodynamic investigation, longitudinal stability during take-offs and landings was satisfactory. Decreasing the area of the hydro-ski 60 percent increased the maximum resistance and emergence speed 40 and 70 percent, respectively. Without the jet exhaust, the resistance was reduced by simulating the vertical-lift component of the forward engines rotated upward. However, the jet exhaust of the forward engines increased the maximum resistance approximately 60 percent. The engine inlets and horizontal tail were free from spray for all loads investigated and for both hydro-ski sizes.

Petynia, William W.↗
NASA NTRS2002

Ground and Flight Evaluation of a Small-Scale Inflatable-Winged Aircraft

A small-scale, instrumented research aircraft was flown to investigate the night characteristics of innersole wings. Ground tests measured the static structural characteristics of the wing at different inflation pressures, and these results compared favorably with analytical predictions. A research-quality instrumentation system was assembled, largely from commercial off-the-shelf components, and installed in the aircraft. Initial flight operations were conducted with a conventional rigid wing having the same dimensions as the inflatable wing. Subsequent flights were conducted with the inflatable wing. Research maneuvers were executed to identify the trim, aerodynamic performance, and longitudinal stability and control characteristics of the vehicle in its different wing configurations. For the angle-of-attack range spanned in this flight program, measured flight data demonstrated that the rigid wing was an effective simulator of the lift-generating capability of the inflatable wing. In-flight inflation of the wing was demonstrated in three flight operations, and measured flight data illustrated the dynamic characteristics during wing inflation and transition to controlled lifting flight. Wing inflation was rapid and the vehicle dynamics during inflation and transition were benign. The resulting angles of attack and of sideslip ere small, and the dynamic response was limited to roll and heave motions.

Murray, James E.↗
NASA NTRS2002

Aerodynamic Database Development for Mars Smart Lander Vehicle Configurations

An aerodynamic database has been generated for the Mars Smart Lander Shelf-All configuration using computational fluid dynamics (CFD) simulations. Three different CFD codes, USM3D and FELISA, based on unstructured grid technology and LAURA, an established and validated structured CFD code, were used. As part of this database development, the results for the Mars continuum were validated with experimental data and comparisons made where applicable. The validation of USM3D and LAURA with the Unitary experimental data, the use of intermediate LAURA check analyses, as well as the validation of FELISA with the Mach 6 CF(sub 4) experimental data provided a higher confidence in the ability for CFD to provide aerodynamic data in order to determine the static trim characteristics for longitudinal stability. The analyses of the noncontinuum regime showed the existence of multiple trim angles of attack that can be unstable or stable trim points. This information is needed to design guidance controller throughout the trajectory.

Bobskill, Glenn J.↗
NASA NTRS2003

Proof Test Diagrams for a Lithia-Alumina-Silica Glass-Ceramic

The glass-ceramic (Zerodur, Schott Glaswerke, Mainz, Germany) contains 70% to 78% by weight crystalline phase of high-quartz structure with a mean crystal size of 50-55 nm. The vitreous phase has a positive thermal expansion coefficient which is practically balanced by the negative coefficient of the crystalline phase. This results in a material which can maintain longitudinal stability during thermal cycling. This was one of the reasons for its choice as the material for the grazing incidence mirrors for the Chandra X-Ray Facility. Brittle materials such as glass and glass-ceramics which exhibit slow crack growth and subsequent fast fracture to failure exhibit a time dependence in strength. The decrease in strength for a constant applied load is known as static fatigue. In many cases, environment plays a major role in the material lifetime. It has been shown for silicate glasses that crack velocity will increase as the amount of water vapor in the environment surface finish and rate of loading. A rough surface finish leads to a lower tensile strength than for an optically polished surface. The strength of glass is observed in general to increase with increasing load rate. This phenomena is known as dynamic fatigue. This was observed for Zerodur by Tucker and Gent and Tucker in previous dynamic fatigue studies, in which lifetimes were obtained. All of the above named factors need to be considered when glass is to be used in load bearing applications.

Tucker, Dennis S.↗
NASA NTRS1961

Low-Speed Investigation of the Effects of Large Wing-Sideslip Angles on the Aerodynamic Characteristics of Two Arrow-Wing-Fuselage Arrangements

A wind-tunnel investigation has been made of two methods proposed to reduce the high sinking speeds and improve the landing characteristics of low-aspect-ratio hypersonic airplanes by placing the wing at large angles of sideslip to increase its effective aspect ratio. The models investigated had conical fuselages and arrow wings, with a leading-edge sweep-back of 77.4deg, an aspect ratio of 1.23, and a 4-percent-thick straight-wedge section. For one model, the wing was pivoted on the fuselage to angles of wing sideslip from 0 to 90deg. For the other model, the wing was fixed to the fuselage, and the wing and fuselage were yawed together to sideslip angles from 0 to 90deg. The investigation was made in the Langley 300-MPH 7- by 10-foot tunnel for an angle-of-attack range from -8 to above 28deg. Longitudinal stability and control through the use of horizontal tails with elevators was studied on the pivoted-wing configuration. The roll control for both configurations was studied with deflection of the apex portion of the wing about an axis along the wing center line. The use of flaps with the wing at large sideslip angles was also investigated.

Thomas G Gainer↗
NASA NTRS1961

Low-Speed Wind-Tunnel Investigation of a Reflection Plane Wing Model Equipped with Partial-Span Jet-Augmented Flaps

An investigation to determine the aerodynamic and flow-field characteristics of three 50-percent partial-span jet-augmented flap configurations located at three spanwise positions has been conducted in the Langley 300 MPH 7- by 10-foot tunnel. The model was a semispan, rectangular unswept wing with a full-span aspect ratio of 8.3 and a thickness-to-chord ratio of 0.167. The results of this investigation showed that an inboard partial-span jet-augmented flap is less effective than a full-span blowing flap. A further reduction in effectiveness is obtained as the blaring is shifted outboard along the span. At a given lift coefficient and angle of attack, the nose-down moment about the quarter chord increases as the blowing is moved outboard. This increase in nose-down moment is due primarily to the fact that the momentum coefficient must be increased in order to obtain the same lift condition. Flow surveys indicate that the tail contribution to static longitudinal stability would be greater for the outboard blowing locations.

Otis, James H., Jr.↗
NASA NTRS1961

Aerodynamic Characteristics in Pitch of a Series of Cruciform-Wing Missiles with Canard Controls 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 static longitudinal stability and control characteristics of a series of missile configurations with canard controls at angles of attack up to about 28 deg. The missiles had cruciform wings and canard surfaces of delta plan form with 70 deg.swept leading edges. Five bodies having fineness ratios of 19.1, 17.7, 16.7, 15.7, and 14.8 were investigated. The results of the investigation indicated a large nonlinear variation of pitching moment with angle of attack for the body of largest fineness ratio that was progressively reduced by decreasing the fineness ratio until it was essentially eliminated for a body of fineness ratio 14.8. The increased linearity of the moment curve would make it possible to reduce the margin of stability so that, for a given canard size and deflection, a higher trim angle of attack might be obtained for the shortest missile than for the longest missile. The pitching-moment results indicated that methods of prediction which assumed linear variations with angle of attack for the wing-alone and wing-plus-interference characteristics were adequate for angles of attack up to about 12 deg. At higher angles of attack it was evident that the characteristics of these components were nonlinear and that more refined methods would be required for adequate prediction.

Spearman, M. Leroy↗
NASA NTRS1961

Aerodynamic Characteristics over a Mach Number Range of 1.40 to 2.78 of a Rocket-Propelled Airplane Configuration having a Low 52.50 Delta Wing and an Unswept Horizontal Tail

A free-flight investigation of an airplane configuration having a low 52.5 deg. delta wing and an unswept horizontal tail has been conducted over a Mach number range of 1.40 to 2.78. At a fixed tail setting of -3.0 deg., the trim lift coefficient and angle of attack varied from about 0.12 to 0.04 and 3.8 deg. to 2.0 deg., respectively. The base drag was approximately 5 percent of the total drag at trim lift. Lift-curve slope, static longitudinal stability, and damping in pitch were obtained only at Mach numbers of 2.59 t o 2.74. Theoretical calculations of lift-curve slope and aerodynamic-center location were in good agreement with experimental results.

Kehlet, Alan B.↗
NASA NTRS2004

Evaluation of a CFD Method for Aerodynamic Database Development using the Hyper-X Stack Configuration

A computational fluid dynamic (CFD) study is performed on the Hyper-X (X-43A) Launch Vehicle stack configuration in support of the aerodynamic database generation in the transonic to hypersonic flow regime. The main aim of the study is the evaluation of a CFD method that can be used to support aerodynamic database development for similar future configurations. The CFD method uses the NASA Langley Research Center developed TetrUSS software, which is based on tetrahedral, unstructured grids. The Navier-Stokes computational method is first evaluated against a set of wind tunnel test data to gain confidence in the code s application to hypersonic Mach number flows. The evaluation includes comparison of the longitudinal stability derivatives on the complete stack configuration (which includes the X-43A/Hyper-X Research Vehicle, the launch vehicle and an adapter connecting the two), detailed surface pressure distributions at selected locations on the stack body and component (rudder, elevons) forces and moments. The CFD method is further used to predict the stack aerodynamic performance at flow conditions where no experimental data is available as well as for component loads for mechanical design and aero-elastic analyses. An excellent match between the computed and the test data over a range of flow conditions provides a computational tool that may be used for future similar hypersonic configurations with confidence.

Parikh, Paresh↗
NASA NTRS1947

Wind-Tunnel Tests of the 1/25-Scale Powered Model of the Martin JRM-1 Airplane. IV - Tests with Ground Board and with Modified Wing and Hull - TED No. NACA 232: Tests with Ground Board and with Modified Wing and Hull, TED No. NACA 232 - Part 4

Wind-tunnel tests were made of a 1/25 scale model of the Martin JRM-1 airplane to determine: (1) The longitudinal stability and control characteristics of the JRM-1 model near the water and lateral and directional stability characteristics with power while moving on the surface of the water, the latter being useful for the design of tip floats; (2) The stability and stalling characteristics of the wing with a modified airfoil contour; (3) Stability characteristics of a hull of larger design gross weight; The test results indicated that the elevator was powerful enough to trim the original model in a landing configuration at any lift coefficient within the specified range of centers of gravity. The ground-board tests for evaluating the aerodynamic forces and moments on an airplane in a simulated cross wind indicate a high dihedral effect in the presence of the ground board and, consequently, during low-speed taxying and take-off, large overturning moments would result which would have to be overcome by the tip floats.

Lockwood, Vernard E.↗