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At least 991 records · Page 55
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

Transonic wing design using potential-flow codes - Successes and failures

The state-of-the-art of transonic wing design by use of computer codes based on the potential flow-theory is presented. The capabilities and limitations of these codes are exemplified by several experiment-theory correlations, including an assessment of pressure distribution from isolated wing and wing-body codes. Computer codes using both conservative and nonconservative differencing schemes were used, and the effects of boundary-layer corrections were considered. Results showed that calculations from a full potential, isolated code correlate well with data from an isolated wing test, but may give poor predictions of the aerodynamic characteristics of some wing-body configurations. Boundary-layer correlations were found to have only moderate effects on experiment-theory correlation. Aeroelastic effects were considered important for high aspect ratio wings of low to moderate thickness, and viscous effects were minimal for typical cruise conditions, even for Reynold's numbers as low as two million. The effect of wind-tunnel walls on experiment-theory correlations remained inconclusive. A wing-body code was used to calculate the flow field about a wing-body configuration with body-mounted engines, typical of the transonic Biz-Jet aircraft.

Hicks, R. M.↗
NASA NTRS1981

Wind-tunnel results for a modified 17-percent-thick low-speed airfoil section

Wind-tunnel tests were conducted in the Langley low-turbulence pressure tunnel to evaluate the effects on performance of modifying a 17-percent-thick low-speed airfoil. The airfoil contour was altered to reduce the pitching-moment coefficient by increasing the forward loading and to increase the climb lift-drag ratio by decreasing the aft upper surface pressure gradient. The tests were conducted over a Mach number range from 0.07 to 0.32, a chord Reynolds number range 1.0 x 10 to the 6th power to 12.0 x 10 to the 6th power, and an angle-of-attack range from about -10 deg to 20 deg.

Mcghee, R. J.↗
NASA NTRS1982

Wind-tunnel investigation of the powered low-speed longitudinal aerodynamics of the Vectored-Engine-Over (VEO) wing fighter configuration

A wind-tunnel investigation incorporating both static and wind-on testing was conducted in the Langley 4- by 7-Meter Tunnel to determine the effects of vectored thrust along with spanwise blowing on the low-speed aerodynamics of an advanced fighter configuration. Data were obtained over a large range of thrust coefficients corresponding to takeoff and landing thrust settings for many nozzle configurations. The complete set of static thrust data and the complete set of longitudinal aerodynamic data obtained in the investigation are presented. These data are intended for reference purposes and, therefore, are presented without analysis or comment. The analysis of the thrust-induced effects found in the investigation are not discussed.

Paulson, J. W.↗
NASA NTRS1982

Performance characteristics of axisymmetric convergent-divergent exhaust nozzles with longitudinal slots in the divergent

An investigation was conducted in the Langley 16 foot Transonic Tunnel and in the static test facility of that tunnel to determine the effects of divergent flap ventilation of an axisymmetric nozzle on nozzle internal (static) and wind on performance. Tests were conducted at 0 deg angle of attack at static conditions and at Mach numbers from 0.6 to 1.2. Ratios of jet total pressure to free stream static pressure were varied from 1.0 (jet off) to approximately 14.0 depending on Mach number. The results of this study indicate that divergent flap ventilation generally provided large performance benefits at overexpanded nozzle conditions and performance reductions at underexpanded nozzle conditions when compared to the baseline (unventilated) nozzles. Ventilation also reduced the peak static and wind on performance levels.

Leavitt, L. D.↗
NASA NTRS1981

A large-scale investigation of engine influence on inlet performance at angle-of-attack

A low-speed wind tunnel test was conducted in the NASA/Ames 40-x-80-foot wind tunnel to investigate the effect of engine/inlet flowfield interaction on inlet performance near flow separation. The effect of engine/inlet flowfield interaction was determined by comparing the performance of a large-scale subsonic inlet (CR = 1.26) close-coupled and remote-coupled to a TF-34 turbofan engine. The remote coupled inlet configuration removes the influence of the engine on the inlet flowfield and further, typifies conventional small-scale inlet test techniques which generally provide no simulation of turbomachinery effects. Test results indicated that engine interaction allows the inlet to operate with lower distortion levels at and beyond the separation angle-of-attack attained without engine interaction.

Hodder, B. K.↗
NASA NTRS1974

Low-speed aerodynamic characteristics of a 42 deg swept high-wing model having a double-slotted flap system and a supercritical airfoil

A low-speed investigation was conducted over an angle-of-attack range from about -4 deg to 20 deg in the Langley V/STOL tunnel to determine the effects of a double-slotted flap, high-lift system on the aerodynamic characteristics of a 42 deg swept high-wing model having a supercritical airfoil. The wing had an aspect ratio of 6.78 and a taper ratio of 0.36; the double-slotted flap consisted of a 35-percent-chord flap with a 15-percent-chord vane. The model was tested with a 15-percent-chord leading-edge slat.

Fournier, P. G.↗
NASA NTRS1973

A Wind-Tunnel Investigation of the Application of the NASA Supercritical Airfoil to a Variable-Wing-Sweep Fighter Airplane

An investigation was conducted in the Langley 8 foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel to evaluate the effectiveness of three variations of the NASA supercritical airfoil as applied to a model of a variable wing sweep fighter airplane. Wing panels incorporating conventional NACA 64A series airfoil with 0.20 and 0.40 camber were used as bases of reference for this evaluation. Static force and moment measurements were obtained for wing leading edge sweep angles of 26, 33, 39, and 72.5 degrees. Fluctuating wing root bending moment data were obtained at subsonic speeds to determine buffet characteristics. Subsonic data were also obtained for determining the effects of wing transition location and spoiler deflection. Limited lateral directional data are included for the conventional 0.20 cambered wing and the supercritical wing.

Ayers, T. G.↗
NASA NTRS1983

Support-sting interference on boattail pressure drag for Reynolds numbers up to 70 x 10 to the 6th

A model was tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel to investigate the effects of Reynolds number on boattail pressure drag for a variety of sting shapes. The boattail pressure drag for constant Mach number increased linearly with Reynolds number over the Reynolds number range tested. The data indicated that, as the disturbance produced by the sting on the boattail increased, the boattail pressure drag became less sensitive to Reynolds number change. Also, it was found that the model base pressure versus Reynolds number curve reached a plateau within the Reynolds number range examined.

Gloss, B. B.↗
NASA NTRS1983

Measurements and predictions of turboprop noise at high cruise speed

Acoustic tests of advanced turboprop (Prop-Fan) models have taken place in or on three facilities: an acoustic wind tunnel, an aerodynamic wind tunnel, and a flight vehicle. Comparisons of data from the three facilities are made for two model designs, the unswept SR-2 and the swept SR-3. The importance of noise propagation through the boundary layer to microphones mounted on the surface of a flight vehicle is discussed and limited comparisons between predictions and measurements are shown. It is concluded that existing methodology is adequate for predictions at high cruise Mach number and that further development is required to improve the methodology for predicting fuselage and wind tunnel boundary layer propagation effects.

Metzger, F. B.↗
NASA NTRS1983

A time-split finite-volume algorithm for three-dimensional flow-field simulation

A general finite-volume algorithm is developed for solving three-dimensional, time-dependent, compressible Navier-Stokes equations for high Reynolds number flows over an arbitrary geometry. This algorithm adapts MacCormack's (1982) explicit-implicit scheme to a time-split, three-dimensional finite-volume concept in a general coordinate system. It is shown that the thin-layer approximation in all three spatial directions significantly reduces the evaluation of viscous terms and allows the algorithm to solve more complicated geometries with all boundaries in two or all three directions. The calculated results using this method are found to be in good agreement with the experimental measurements of a blunt-fin induced shock wave and boundary-layer interaction problems. Observations of the existence of peak pressure, primary horseshoe and secondary vortices, and reversed supersonic zones show that computational fluid dynamics can effectively supplement the wind tunnel tests for aerodynamic design as well as for understanding basic fluid dynamics.

Hung, C. M.↗
NASA NTRS1984

Computer studies of hybrid-slotted working sections with minimum interference at subsonic speeds

A series of computations on tunnel boundary-interference effects for hybrid-slotted working sections was performed using the WALINT code. The slots were modeled as lines of porosity with linear crossflow characteristics. The basic shape evaluated was for a rectangular section with height-to-width ratio = 0.835 and its companion in the duplex mode (half model testing) with height-to-width ratio = 0.6. A best overall basic configuration was determined with seven slots on each wall with open area ratio on each wall of 17.5%. For both full-span and half-model testing, the optimum solution required closing all but two slots on each of the half-walls parallel to the plane of the wing (equivalent to four slots on the full floor and ceiling). The results are presented here for the best configurations and are shown to be within the figure-of-merit range of + or - 0.04 in upwash, and + or - 0.1 in curvature for the Mach number range 0.6 to 0.85. Blockage effects are shown to be small.

Steinle, F. W., Jr.↗
NASA NTRS1984

Stability and control results for advanced turboprop aft-mount installations

An experimental investigation was recently conducted in the NASA-Langley Research Center 4- by 7-Meter Tunnel to explore the effects of aft-mounted advanced turboprop installations on the low-speed stability and control characteristics of a representative transport aircraft configuration. For the conditions investigated, the experimental results indicate that the longitudinal and directional stability of the configuration was not adversely affected by aft-pylon-mounted single-rotation tractor or counter-rotation pusher propeller installations. The results indicate that for the single-rotation tractor configuration the propeller induced aerodynamics are dominated by the interaction of the propeller slipsteam with the pylon, and that the propeller contribution to the stability characteristics for the counter-rotating pusher configuration is dominated by propeller normal forces. The longitudinal and directional control effectiveness, engine-out characteristics, and ground effects are also presented.

Coe, P. L., Jr.↗
NASA NTRS1985

Aerodynamic characteristics of a high-wing transport configuration with a over-the-wing nacelle-pylon arrangement

An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects on the aerodynamic characteristics of a high-wing transport configuration of installing an over-the-wing nacelle-pylon arrangement. The tests are conducted at Mach numbers from 0.70 to 0.82 and at angles of attack from -2 deg to 4 deg. The configurational variables under study include symmetrical and contoured nacelles and pylons, pylon size, and wing leading-edge extensions. The symmetrical nacelles and pylons reduce the lift coefficient, increase the drag coefficient, and cause a nose-up pitching-moment coefficient. The contoured nacelles significantly reduce the interference drag, though it is still excessive. Increasing the pylon size reduces the drag, whereas adding wing leading-edge extension does not affect the aerodynamic characteristics significantly.

Henderson, W. P.↗
NASA NTRS1986

The application to airfoils of a technique for reducing orifice-induced pressure error at high Reynolds numbers

A wind tunnel investigation was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel to study the effects of porous (sintered metal) plug orifices on orifice-induced static-pressure measurement error at high Reynolds numbers. A NACA airfoil was tested at Mach numbers from 0.60 to 0.80 and at Reynolds numbers from 6 x 1,000,000 to 40 x 1,000,000. Data are included which compare pressure measurements obtained from porous plug orifices and from conventional orifices with diameters of 0.025 cm (0.010 in.) and 0.102 cm (0.040 in.). The two dimensional airfoil code GRUMFOIL was used to calculate boundary layer displacement thickness. The response time and the downstream effect of the porous plug orifice were considered in this investigation. The results showed that the porous plug orifice could be a viable method of reducing pressure error. The data also showed that the pressure measurements obtained with a 0.102-cm-diameter orifice were very close to the measurements obtained with 0.025-cm-diameter orifice over such of the airfoil and that downstream of a shock the orifice size was not critical.

Plentovich, E. B.↗
NASA NTRS1986

Low-speed stability and control characteristics of a transport model with aft-fuselage-mounted advanced turboprops

A limited experimental investigation was conducted in the Langley 4- by 7-Meter Tunnel to explore the effects of aft-fuselage-mounted advanced turboprop installations on the low-speed stability and control characteristics of a representative transport aircraft in a landing configuration. In general, the experimental results indicate that the longitudinal and lateral-directional stability characteristics for the aft-fuselage-mounted single-rotation tractor and counter-rotation pusher propeller configurations tested during this investigation are acceptable aerodynamically. For the single-rotation tractor configuration, the propeller-induced aerodynamics are significantly influenced by the interaction of the propeller slipstream with the pylon and nacelle. The stability characteristics for the counter-rotation pusher configuration are strongly influenced by propeller normal forces. The longitudinal and directional control effectiveness, engine-out characteristics, and ground effects are also presented. In addition, a tabulated presentation of all aerodynamic data presented in this report is included as an appendix.

Applin, Z. T.↗
NASA NTRS1986

Simulation investigation of the effect of the NASA Ames 80-by 120-foot wind tunnel exhaust flow on light aircraft operating in the Moffett field trafffic pattern

A preliminary study of the exhaust flow from the Ames Research Center 80 by 120 Foot Wind Tunnel indicated that the flow might pose a hazard to low-flying light aircraft operating in the Moffett Field traffic pattern. A more extensive evaluation of the potential hazard was undertaken using a fixed-base, piloted simulation of a light, twin-engine, general-aviation aircraft. The simulated aircraft was flown through a model of the wind tunnel exhaust by pilots of varying experience levels to develop a data base of aircraft and pilot reactions. It is shown that a light aircraft would be subjected to a severe disturbance which, depending upon entry condition and pilot reaction, could result in a low-altitude stall or cause damage to the aircraft tail structure.

Streeter, Barry G.↗
NASA NTRS1987

An experimental investigation of an advanced turboprop installation on a swept wing at subsonic and transonic speeds

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of a turboprop-nacelle installation on the pressure distributions over a swept, supercritical wing. The tests were conducted at Mach numbers from 0.20 to 0.80, at angles of attack from 0 to 5 degrees, nacelle nozzle pressure ratios from 1.0 to 1.6, and at propeller tip speeds from 700 to 800 ft/sec. The results of this study indicate that the turboprop nacelle interference, with and without power, on a swept wing is greater on the inboard wing panel than on the outboard wing panel. The over-the-wing nacelle installation with the propeller upwash on the inboard panel had flow separation problems at a Mach number of 0.80. No severe flow separation problems appear to exist for either propeller rotation direction for the under-the-wing nacelle installation. The local flow disturbances caused by the under-the-wing nacelle installation were in general less severe than for the over-the-wing nacelle installation.

Carlson, John R.↗
NASA NTRS1987

Comparison of wind tunnel and flight test afterbody and nozzle pressures for a twin-jet fighter aircraft at transonic speeds

Afterbody and nozzle pressures measured on a 1/12-scale model and in flight on a twin-jet fighter aircraft were compared as Mach number varied from 0.6 to 1.2, Reynolds number from 17.5 million to 302.5 million, and angle of attack from 1 to 7 deg. At Mach 0.6 and 0.8, nozzle pressure coefficient distributions and nozzle axial force coefficients agreed and showed good recompression. At Mach 0.9 and 1.2, flow complexity caused a loss in recompression for both flight and wind tunnel nozzle data. The flight data exhibited less negative values of pressure coefficient and lower axial force coefficients than did the wind tunnel data. Reynolds number effects were noted only at these Mach numbers. Jet temperature and mass flux ratio did not affect the comparisons of nozzle axial flow coefficient. At subsonic speeds, the levels of pressure coefficient distributions on the upper fuselage and lower nacelle surfaces for flight were less negative than those for the model. The model boundary layer thickness at the aft rake station exceeded that for the forward rake station and increased with increasing angle of attack. The flight boundary layer thickness at the aft rake station was less than that for the forward rake station and decreased with increasing angle of attack.

Nugent, Jack↗