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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Lift augmentation on a delta wing via leading edge fences and the Gurney flap

Wind tunnel tests were conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

Buchholz, Mark D.↗
NASA NTRS1993

Lift Augmentation on a Delta Wing via Leading Edge Fences and the Gurney Flap

Wind tunnel tests have been conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve by as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

Buchholz, Mark D.↗
NASA NTRS1972

Flight measurements of buffet characteristics of the F-104 airplane for selected wing-flap deflections

A flight program was conducted on the F-104 airplane to investigate the effects of moderate deflections of wing leading- and trailing-edge flaps on the buffet characteristics at subsonic and transonic Mach numbers. Selected deflections of the wing leading and trailing-edge flaps, individually and in combination, were used to assess buffet onset, intensity, and frequency; lift curves; and wing-rock characteristics for each configuration. Increased deflection of the trailing-edge flap delayed the buffet onset and buffet intensity rise to a significantly higher airplane normal-force coefficient. Deflection of the leading-edge flap produced some delay in buffet onset and the resulting intensity rise at low subsonic speeds. Increased deflection of the trailing-edge flap provided appreciable lift increments in the angle-of-attack range covered, whereas the leading-edge flap provided lift increments only at high angles-of-attack. The pilots appreciated the increased maneuvering envelope provided by the flaps because of the improved turn capability.

Friend, E. L.↗
NASA NTRS1984

Performance degradation of a typical twin engine commuter type aircraft in measured natural icing conditions

The performance of an aircraft in various measured icing conditions was investigated. Icing parameters such as liquid water content, temperature, cloud droplet sizes and distributions were measured continuously while in icing. Flight data were reduced to provide plots of the aircraft drag polars and lift curves (CL vs. alpha) for the measured ""iced'' condition as referenced to the uniced aircraft. These data were also reduced to provide plots of thrust horsepower required vs. single engine power available to show how icing affects engine out capability. It is found that performance degradation is primarily influenced by the amount and shape of the accumulated ice. Glaze icing caused the greatest aerodynamic performance penalties in terms of increased drag and reduction in lift while aerodynamic penalties due to rime icing were significantly lower.

Ranaudo, R. J.↗
NASA NTRS1984

Performance degradation of a typical twin engine commuter type aircraft in measured natural icing conditions

The performance of an aircraft in various measured icing conditions was investigated. Icing parameters such as liquid water content, temperature, cloud droplet sizes and distributions were measured continuously while in icinig. Flight data wre reduced to provide plots of the aircraft drag polars and lift curves (CL vs. alpha) for the measured 'iced' condition as referenced to the uniced aircraft. These data were also reduced to provide plots of thrust horsepower required vs. single engine power available to show how icing affects engine out capability. It is found that performance degradation is primarily influenced by the amount and shape of the accumulated ice. Glaze icing caused the greatest aerodynamic performance penalties in terms of increased drag and reduction in lift while aerodynamic penalties due to rime icing were significantly lower. Previously announced in STAR as N84-13173

Ranaudo, R. J.↗
NASA NTRS2014

High-Lift OVERFLOW Analysis of the DLR-F11 Wind Tunnel Model

In response to the 2nd AIAA CFD High Lift Prediction Workshop, the DLR-F11 wind tunnel model is analyzed using the Reynolds-averaged Navier-Stokes flow solver OVERFLOW. A series of overset grids for a bracket-off landing configuration is constructed and analyzed as part of a general grid refinement study. This high Reynolds number (15.1 million) analysis is done at multiple angles-of-attack to evaluate grid resolution effects at operational lift levels as well as near stall. A quadratic constitutive relation recently added to OVERFLOW for improved solution accuracy is utilized for side-of-body separation issues at low angles-of-attack and outboard wing separation at stall angles. The outboard wing separation occurs when the slat brackets are added to the landing configuration and is a source of discrepancy between the predictions and experimental data. A detailed flow field analysis is performed at low Reynolds number (1.35 million) after pressure tube bundles are added to the bracket-on medium grid system with the intent of better understanding bracket/bundle wake interaction with the wing's boundary layer. Localized grid refinement behind each slat bracket and pressure tube bundle coupled with a time accurate analysis are exercised in an attempt to improve stall prediction capability. The results are inconclusive and suggest the simulation is missing a key element such as boundary layer transition. The computed lift curve is under-predicted through the linear range and over-predicted near stall, and the solution from the most complete configuration analyzed shows outboard wing separation occurring behind slat bracket 6 where the experiment shows it behind bracket 5. These results are consistent with most other participants of this workshop.

High-lift↗
NASA NTRS1934

Effect of the Surface Condition of a Wing on the Aerodynamic Characteristics of an Airplane

In order to determine the effect of the surface conditions of a wing on the aerodynamic characteristics of an airplane, tests were conducted in the N.A.C.A. full-scale wind tunnel on the Fairchild F-22 airplane first with normal commercial finish of wing surface and later with the same wing polished. Comparison of the characteristics of the airplane with the two surface conditions shows that the polish caused a negligible change in the lift curve, but reduced the minimum drag coefficient by 0.001. This reduction in drag if applied to an airplane with a given speed of 200 miles per hour and a minimum drag coefficient of 0.025 would increase the speed only 2.9 miles per hour, but if the speed remained the same, the power would be reduced 4 percent.

AIRPLANES - FAIRCHILD 22 (Amer ) WIND TUNNEL TEST↗
NASA NTRS1980

Long duct nacelle aerodynamic development for DC-10 derivatives

The results are presented of a wind tunnel test utilizing a 4.7-percent-scale semispan model of the DC-10 in the Calspan 8-foot transonic wind tunnel. The effect of a revised long-duct nacelle shape on the channel velocities, the incremental drag relative to the baseline long-duct nacelle, and channel velocities for the baseline long-duct nacelle were determined and compared with data obtained at Ames. The baseline and the revised long-duct nacelles are representative of a CF6-50 mixed-flow configuration and were evaluated on a model of a proposed DC-10 stretched-fuselage configuration. The results showed that the revised long-duct nacelle has an appreciable effect on the inboard channel velocities, resulting in an increased channel Mach number. However, the pressure recovery on the nacelle afterbody was about the same for both nacelles. The lift curves for both long-duct nacelle configurations were the same. The channel pressures measured at Calspan were in good agreement with those measured at Ames for the baseline long-duct nacelle. The incremental drag for the revised nacelle was measured as two to four counts (three counts is approximately equal to one percent of the airplane drag) higher than that of the baseline long-duct nacelle.

Patel, S. P.↗
NASA NTRS1983

An experimental investigation of the low Reynolds number performance of the Lissaman 7769 airfoil

A Lissaman 7769 airfoil, used on the Gossamer Condor and Gossamer Albatross human-powered aircraft, was tested in a low turbulence subsonic wind tunnel. Lift and drag data were collected at chord Reynolds numbers of 100,000, 150,000, 200,000, 250,000, and 300,000; at angles of attack from -10 to +20 deg by using an external strain gage force balance. Lift curves, drag curves, and drag polars were generated from both uncorrected data and data corrected for wind tunnel blockage effects. A flow visualization study was performed to correlate with the force data. The results of the investigation have shown that the airfoil exhibits a significant degradation in performance for chord Reynolds numbers below 150,000.

Conigliaro, P. E.↗
NASA NTRS1987

Application of an upwind Navier-Stokes code to two-dimensional transonic airfoil flow

An upwind-biased implicit approximate factorization Navier-Stokes algorithm is applied to a variety of steady transonic airfoil cases, using the NACA 0012, RAE 2822, and Jones supercritical airfoils. The thin-layer form of the compressible Navier-Stokes equations is used. Both the CYBER 205 and CRAY 2 supercomputers are utilized, with average computational speeds of about 18 and 16 microsec/gridpoint/iteration, respectively. Lift curves, drag polars, and variations in drag coefficient with Mach number are determined for the NACA 0012 and Jones supercritical airfoils. Also, several cases are computed for comparison with experiment. The effect of grid density and grid extent on a typical turbulent airfoil solution is shown. An algebraic eddy-viscosity turbulence model is used for all of the computations.

Rumsey, Christopher L.↗
NASA NTRS1988

Development of a real-time aeroperformance analysis technique for the X-29A advanced technology demonstrator

The X-29A advanced technology demonstrator has shown the practicality and advantages of the capability to compute and display, in real time, aeroperformance flight results. This capability includes the calculation of the in-flight measured drag polar, lift curve, and aircraft specific excess power. From these elements many other types of aeroperformance measurements can be computed and analyzed. The technique can be used to give an immediate postmaneuver assessment of data quality and maneuver technique, thus increasing the productivity of a flight program. A key element of this new method was the concurrent development of a real-time in-flight net thrust algorithm, based on the simplified gross thrust method. This net thrust algorithm allows for the direct calculation of total aircraft drag.

Ray, R. J.↗
NASA NTRS1988

Development of a real-time aeroperformance analysis technique for the X-29A advanced technology demonstrator

The X-29A advanced technology demonstrator has shown the practicality and advantages of the capability to compute and display, in real time, aeroperformance flight results. This capability includes the calculation of the in flight measured drag polar, lift curve, and aircraft specific excess power. From these elements, many other types of aeroperformance measurements can be computed and analyzed. The technique can be used to give an immediate postmaneuver assessment of data quality and maneuver technique, thus increasing the productivity of a flight program. A key element of this new method was the concurrent development of a real-time in flight net thrust algorithm, based on the simplified gross thrust method. This net thrust algorithm allows for the direct calculation of total aircraft drag.

Ray, R. J.↗
NASA NTRS1989

Capabilities of wind tunnels with two-adaptive walls to minimize boundary interference in 3-D model testing

An initial wind tunnel test was made to validate a new wall adaptation method for 3-D models in test sections with two adaptive walls. First part of the adaptation strategy is an on-line assessment of wall interference at the model position. The wall induced blockage was very small at all test conditions. Lift interference occurred at higher angles of attack with the walls set aerodynamically straight. The adaptation of the top and bottom tunnel walls is aimed at achieving a correctable flow condition. The blockage was virtually zero throughout the wing planform after the wall adjustment. The lift curve measured with the walls adapted agreed very well with interference free data for Mach 0.7, regardless of the vertical position of the wing in the test section. The 2-D wall adaptation can significantly improve the correctability of 3-D model data. Nevertheless, residual spanwise variations of wall interference are inevitable.

Rebstock, Rainer↗
NASA NTRS1991

Aerodynamic parameter studies and sensitivity analysis for rotor blades in axial flight

The analytical capability is offered for aerodynamic parametric studies and sensitivity analyses of rotary wings in axial flight by using a 3-D undistorted wake model in curved lifting line theory. The governing equations are solved by both the Multhopp Interpolation technique and the Vortex Lattice method. The singularity from the bound vortices is eliminated through the Hadamard's finite part concept. Good numerical agreement between both analytical methods and finite differences methods are found. Parametric studies were made to assess the effects of several shape variables on aerodynamic loads. It is found, e.g., that a rotor blade with out-of-plane and inplane curvature can theoretically increase lift in the inboard and outboard regions respectively without introducing an additional induced drag.

Chiu, Y. Danny↗
NASA NTRS1991

Aerodynamic parametric studies and sensitivity analysis for rotor blades in axial flight

The analytical capability is offered for aerodynamic parametric studies and sensitivity analyses of rotary wings in axial flight by using a 3D undistorted wake model in curved lifting line theory. The governing equations are solved by both the Multhopp Interpolation technique and the Vortex Lattice method. The singularity from the bound vortices is eliminated through the Hadamard's finite part concept. Good numerical agreement between both analytical methods and finite differences methods are found. Parametric studies were made to assess the effects of several shape variables on aerodynamic loads. It is found, e.g., that a rotor blade with out-of-plane and inplane curvature can theoretically increase lift in the inboard and outboard regions respectively without introducing an additional induced drag.

Chiu, Y. D.↗
NASA NTRS1930

Tests of Large Airfoils in the Propeller Research Tunnel, Including Two with Corrugated Surfaces

This report gives the results of the tests of seven 2 by 12 foot airfoils (Clark Y, smooth and corrugated, Gottingen 398, N.A.C.A. M-6, and N.A.C.A. 84). The tests were made in the propeller research tunnel of the National Advisory Committee for Aeronautics at Reynolds numbers up to 2,000,000. The Clark Y airfoil was tested with three degrees of surface smoothness. Corrugating the surface causes a flattening of the lift curve at the burble point and an increase in drag at small flying angles.

Wood, Donald H↗
NASA NTRS1999

Investigation of a Technique for Measuring Dynamic Ground Effect in a Subsonic Wind Tunnel

To better understand the ground effect encountered by slender wing supersonic transport aircraft, a test was conducted at NASA Langley Research Center's 14 x 22 foot Subsonic Wind Tunnel in October, 1997. Emphasis was placed on improving the accuracy of the ground effect data by using a "dynamic" technique in which the model's vertical motion was varied automatically during wind-on testing. This report describes and evaluates different aspects of the dynamic method utilized for obtaining ground effect data in this test. The method for acquiring and processing time data from a dynamic ground effect wind tunnel test is outlined with details of the overall data acquisition system and software used for the data analysis. The removal of inertial loads due to sting motion and the support dynamics in the balance force and moment data measurements of the aerodynamic forces on the model is described. An evaluation of the results identifies problem areas providing recommendations for future experiments. Test results are validated by comparing test data for an elliptical wing planform with an Elliptical wing planform section with a NACA 0012 airfoil to results found in current literature. Major aerodynamic forces acting on the model in terms of lift curves for determining ground effect are presented. Comparisons of flight and wind tunnel data for the TU-144 are presented.

Graves, Sharon S.↗
NASA NTRS1976

Reduction of Stall-Spin Entry Tendencies Through Wing Aerodynamic Design

Results of recent wind tunnel tests and analyses are described which are aimed at helping to prevent spin entry after a stall by modifications to the aircraft wing. Simple add-on devices were utilized which modify the shape of the lift curve after stall, making it essentially flat to 30- 40 degree angle of attack. It is hypothesized that his will greatly reduce the danger of spin entry. Further research will be aimed at laying a firmer theoretical and analytical basis for the approach and at incorporating the complete full scale airplane and its dynamics into the problem.

Kroeger, R. A.↗