Search NASA⌕ Search

SEARCH · Search NASA

Results for “MACH NUMBER EFFECTS - WING SECTIONS”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Impact probe displacement in a supersonic turbulent boundary layer.

The experiment was performed on the test section sidewall in a supersonic pressure tunnel. The boundary layer at the test station was surveyed in turn by each of 8 impact probes ranging in size from about 1.3 to 48 mm. The impact pressures measured by these probes were combined with the test section static pressure to calculate Mach numbers. Probe displacement effects were evaluated in terms of these Mach number values.

Allen, J. M.↗
NASA NTRS1980

Analytical investigation of critical phenomena in MHD power generators

Critical phenomena in the Arnold Engineering Development Center (AEDC) High Performance Demonstration Experiment (HPDE) and the U.S. U-25 Experiment, are analyzed. The performance of a NASA specified 500 MW(th) flow train is analyzed. Critical phenomena analyzed include: Hall voltage overshoots; optimal load schedules; parametric dependence of the electrode voltage drops; boundary layer behavior; near electrode phenomena with finite electrode segmentation; current distribution in the end regions; scale up rules; optimum Mach number distribution; and the effects of alternative cross sectional shapes.

Source record↗
NASA NTRS2000

Airfoil Dynamic Stall and Rotorcraft Maneuverability

The loading of an airfoil during dynamic stall is examined in terms of the augmented lift and the associated penalties in pitching moment and drag. It is shown that once stall occurs and a leading-edge vortex is shed from the airfoil there is a unique relationship between the augmented lift, the negative pitching moment, and the increase in drag. This relationship, referred to here as the dynamic stall function, shows limited sensitivity to effects such as the airfoil section profile and Mach number, and appears to be independent of such parameters as Reynolds number, reduced frequency, and blade sweep. For single-element airfoils there is little that can be done to improve rotorcraft maneuverability except to provide good static C(l(max)) characteristics and the chord or blade number that is required to provide the necessary rotor thrust. However, multi-element airfoils or airfoils with variable geometry features can provide augmented lift in some cases that exceeds that available from a single-element airfoil. The dynamic stall function is shown to be a useful tool for the evaluation of both measured and calculated dynamic stall characteristics of single element, multi-element, and variable geometry airfoils.

Bousman, William G.↗
NASA NTRS1986

Experimental evaluation of wall Mach number distributions of the octagonal test section proposed for NASA Lewis Research Center's altitude wind tunnel

Wall Mach number distributions were determined over a range of test-section free-stream Mach numbers from 0.2 to 0.92. The test section was slotted and had a nominal porosity of 11 percent. Reentry flaps located at the test-section exit were varied from 0 (fully closed) to 9 (fully open) degrees. Flow was bled through the test-section slots by means of a plenum evacuation system (PES) and varied from 0 to 3 percent of tunnel flow. Variations in reentry flap angle or PES flow rate had little or no effect on the Mach number distributions in the first 70 percent of the test section. However, in the aft region of the test section, flap angle and PES flow rate had a major impact on the Mach number distributions. Optimum PES flow rates were nominally 2 to 2.5 percent wtih the flaps fully closed and less than 1 percent when the flaps were fully open. The standard deviation of the test-section wall Mach numbers at the optimum PES flow rates was 0.003 or less.

Harrington, Douglas E.↗
NASA NTRS1947

The Effectiveness at High Speeds of a 20-Percent-chord Plain Trailing-edge Flap on the NACA 65-210 Airfoil Section

An analysis has been made of the lift-control effectiveness of a 20-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section from section lift-coefficient data obtained at Mach numbers from 0.3 to 0.875. In addition, the effectiveness of the plain flap as a lift-control device has been compared with the corresponding effectiveness of both a spoiler and a dive-recovery flag on the INCA 65-210 airfoil section.

Stivers, Louis S., Jr.↗
NASA NTRS1971

Aerodynamic characteristics of an all-body hypersonic aircraft configuration at Mach numbers from 0.65 to 10.6

Aerodynamic characteristics of a model designed to represent an all body, hypersonic cruise aircraft are presented for Mach numbers from 0.65 to 10.6. The configuration had a delta planform with an elliptic cone forebody and an afterbody of elliptic cross section. Detailed effects of varying angle of attack (-2 to +15 deg), angle of sideslip (-2 to +8 deg), Mach number, and configuration buildup were considered. In addition, the effectiveness of horizontal tail, vertical tail, and canard stabilizing and control surfaces was investigated. The results indicate that all configurations were longitudinally stable near maximum lift drag ratio. The configurations with vertical tails were directionally stable at all angles of attack. Trim penalties were small at hypersonic speeds for a center of gravity location representative of the airplane, but because of the large rearward travel of the aerodynamic center, trim penalties were severe at transonic Mach numbers.

Nelms, W. P., Jr.↗
NASA NTRS1974

Calibration of the Langley 16-foot transonic tunnel with test section air removal

The Langley 16-foot transonic tunnel with test section air removal (plenum suction) was calibrated to a Mach number of 1.3. The results of the calibration, including the effects of slot shape modifications, test section wall divergence, and water vapor condensation, are presented. A complete description of the wind tunnel and its auxiliary equipment is included.

Corson, B. W., Jr.↗
NASA NTRS1959

Experimental Investigation of Expanded Duct Sections and Screens for Reducing FLow Distortions at Subsonic Flows

An investigation of expanded duct sections and the effect of their design parameters on flow distortion over a duct Mach number range of 0.19 to 0.67 was conducted in the small tunnel facility of the Lewis Research Center. The parameters investigated were: (1) entrance angle of expanded section, (2) length of expanded section, (3) area ratio of expanded section, (4) location of expanded section relative to the engine face, and (5) the use of screens of varying solidities and mesh. Expansion half-angles of deg, 15 deg, and 30 deg reduced the total-pressure distortions induced in the duct. The larger expansion angles reduced circumferential distortion more effectively than radial distortion. However, the half-angle of 15 deg appeared to be optimum for reducing both radial and circumferential distortions while still maintaining a high total-pressure recovery. Increasing the expanded-section area ratio and increasing the expanded-section lengths with-the 150 expansion half-angle led to less total-pressure distortion with no appreciable loss in pressure recovery. Screens incorporated in the expanded section indicated that 22.2-percent- solidity screens decreased distortion still further.while 37.3-percent- solidity screens generally increased distortion above that of a constant- area duct incorporating the same solidity screen.

Chiccine, Bruce G.↗
NASA NTRS1979

Effect of nose bluntness and afterbody shape on aerodynamic characteristics of a monoplanar missile concept with bodies of circular and elliptical cross sections at a Mach number of 2.50

The tests were performed at a Mach number of 2.50 and at angles of attack from about -4 deg to 32 deg. The results indicate that increasing nose bluntness increases zero lift drag and decreases both the maximum lift-drag ratio and the level of directional stability. The center of pressure generally moves forward with increasing nose size; however, small nose radii on the modified elliptical configurations move the center of pressure rearward. The circular bodied configurations exhibit the greatest longitudinal stability and the least directional stability. Concepts with the variable geometry afterbody contour display the most directional stability and the greatest zero lift drag.

Graves, E. B.↗
NASA NTRS1987

Experimental evaluation of blockage ratio and plenum evacuation system flow effects on pressure distribution for bodies of revolution in 0.1 scale model test section of NASA Lewis Research Center's proposed altitude wind tunnel

An experimental investigation was conducted in the slotted test section of the 0.1-scale model of the proposed Altitude Wind Tunnel to evaluate wall interference effects at tunnel Mach numbers from 0.70 to 0.95 on bodies of revolution with blockage rates of 0.43, 3, 6, and 12 percent. The amount of flow that had to be removed from the plenum chamber (which surrounded the slotted test section) by the plenum evacuation system (PES) to eliminate wall interference effects was determined. The effectiveness of tunnel reentry flaps in removing flow from the plenum chamber was examined. The 0.43-percent blockage model was the only one free of wall interference effects with no PES flow. Surface pressures on the forward part of the other models were greater than interference-free results and were not influenced by PES flow. Interference-free results were achieved on the aft part of the 3- and 6-percent blockage models with the proper amount of PES flow. The required PES flow was substantially reduced by opening the reentry flaps.

Burley, Richard R.↗