Search NASA⌕ Search

SEARCH · Search NASA

Results for “MACH NUMBER EFFECTS - PROPELLERS”

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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Investigation of the asymmetric aerodynamic characteristics of cylindrical bodies of revolution with variations in nose geometry and rotational orientation at angles of attack to 58 degrees and Mach numbers to 2

Wind-tunnel tests were conducted to investigate the side forces and yawing moments that can occur at high angles of attack and zero sideslip for cylindrical bodies of revolution. Two bodies having several tangent ogive forebodies with fineness ratios of 0.5, 1.5, 2.5, and 3.5 were tested. The forebodies with fineness ratios of 2.5 and 3.5 had several bluntnesses. The cylindrical afterbodies had fineness ratios of 7 and 13. The model components - tip, forebody, and afterbody - were tested in various rotational positions about their axes of symmetry. Most of the tests were conducted at a Mach number of 0.25, a Reynolds number of 0.32 x 10 to the 6th power, and with the afterbody that had a fineness ratio of 7 and with selected forebodies. The effect of Mach number was determined with the afterbody that had a fineness ratio of 13 and with selected forebodies at mach numbers from 0.25 to 2 at Reynolds number = 0.32 X 10 to the 6th power. Maximum angle of attack was 58 deg.

Kruse, R. L.↗
NASA NTRS1991

Waves and thermodynamics in high Mach number propulsive ducts

The propulsive effects of waves in ducts, especially at high Mach numbers, are investigated, focusing on drag and thrust and on the conversion of heat into waves which produce thrust. It is shown that essentially all of the work done by an expanding fluid passing through a duct at high Mach number is delivered in the form of waves, and that duct surface angles exist that are optimum for the production of thrust from a wave. The effects of wave phenomena on drag and thrust are considered by extending the concept of a Busemann biplane into that of a 'Busemann scramjet, taking 'off-design' performance into account. An idealized model of a streamtube with heat addition is developed, and flow mechanisms involved in generating thrust by the expansion of this streamtube in an exhaust nozzle are examined.

Stalker, R. J.↗
NASA NTRS1972

Assessment of a transitional boundary layer theory at low hypersonic Mach numbers

An investigation was carried out to assess the accuracy of a transitional boundary layer theory in the low hypersonic Mach number regime. The theory is based upon the simultaneous numerical solution of the boundary layer partial differential equations for the mean motion and an integral form of the turbulence kinetic energy equation which controls the magnitude and development of the Reynolds stress. Comparisions with experimental data show the theory is capable of accurately predicting heat transfer and velocity profiles through the transitional regime and correctly predicts the effects of Mach number and wall cooling on transition Reynolds number. The procedure shows promise of predicting the initiation of transition for given free stream disturbance levels. The effects on transition predictions of the pressure dilitation term and of direct absorption of acoustic energy by the boundary layer were evaluated.

Shamroth, S. J.↗
NASA NTRS1960

A Note on the Drag Due to Lift of Delta Wings at Mach Numbers up to 2.0

In order to indicate the effects of Reynolds number and other variables on the drag due to lift of delta wings for Mach numbers up to 2.0, the results of several investigations of wing-body combinations having plane delta wings with aspect ratios from 2 to 4 have been assembled for comparison and brief analysis. The effects of Reynolds number, leading-edge radius, and thickness ratio could generally be correlated with Reynolds number based on the leading-edge radius as a parameter. The effects of leading-edge Reynolds number on drag due to lift were large at Mach numbers less than 0.25. However, with increases in Mach number, the effects decreased and were almost negligible at a Mach number of 2.0. and trimming were large, as would be expected. The effects of aspect ratio and trimming were large, as would be expected. It was indicated at least for subsonic and transonic speeds that improvement in the drag due to lift might be obtained from wing modifications designed to inhibit flow separation.

Robert S. Osborne↗
NASA NTRS1959

Effect of External Boundary Layer on Performance of Axisymmetric Inlet at Mach Numbers of 3.0 and 2.5

The effect of an external boundary layer on the performance of an axisymmetric external-internal-compression inlet was evaluated at Mach numbers of 3.0 and 2.5 and Reynolds numbers from 2.2 to 0.5 x 10(exp 6) per foot. The inlet was tested at locations up to two-thirds of the way into the 1.7- and 9.0-inch boundary layers generated by a flat plate and the tunnel floor, respectively. The inlet could be readily started at all conditions tested, including those where the boundary layer was separated upstream of the inlet by the various shock systems during the restart cycle. Although the inlet performance decreased with increasing immersion into the boundary layer at both Mach numbers, the inlet was more sensitive to boundary-layer ingestion at the design Mach number of 3.0.

Samanich, N. E.↗
NASA NTRS1973

Pressure signatures for a .00053 scale model of the Saturn 5-Apollo launch vehicle with simulated exhaust plumes

Wind tunnel pressure signatures are presented for 10 Mach numbers over a range from 3.01 to 7.29 for a .00053-scale model of the Saturn 5 Apollo launch vehicle complete with escape tower and solid body simulated exhaust plumes for each Mach number. The effect of simulated plume length on the wind tunnel pressure signature was investigated at Mach 4.01. An analysis of the error incurred by extrapolating pressure signatures having strong shock waves by weak shock procedures is presented.

Hicks, R. M.↗
NASA NTRS1985

Effects of empennage surface location on aerodynamic characteristics of a twin-engine afterbody model with nonaxisymmetric nozzles

An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of empennage surface location and vertical tail cant angle on the aft-end aerodynamic characteristics of a twin-engine fighter-type configuration. The configuration featured two-dimensional convergent-divergent nozzles and twin-vertical tails. The investigation was conducted with different empennage locations that included two horizontal and three vertical tail positions. Vertical tail cant angle was varied from -10 deg to 20 deg for one selected configuration. Tests were conducted at Mach number 0.60 to 1.20 and at angles of attack -3 to 9 deg. Nozzle pressure ratio was varied from jet off to approximately 9, depending upon Mach number. Tail interference effects were present throughout the range of Mach numbers tested and found to be either favorable or adverse, depending upon test condition and model configuration. At a Mach number of 0.90, adverse interference effects accounted for a significant percentage of total aft-end drag. Interference effects on the nozzle were generally favorable but became adverse as the horizontal tails were moved from a mid to an aft position. The configuration with nonaxisymmetric nozzles had lower total aft-end drag with tails-off than a similar configuration with axisymmetric nozzles at Mach numbers of 0.60 and 0.90.

Capone, Francis J.↗
NASA NTRS1975

Wind tunnel investigation of the aerodynamic characteristics of five forebody models at high angles of attack at Mach numbers from 0.25 to 2

Five forebody models of various shapes were tested in the Ames 6- by 6-Foot Wind Tunnel to determine the aerodynamic characteristics at Mach numbers from 0.25 to 2 at a Reynolds number of 800000. At a Mach number of 0.6 the Reynolds number was varied from 0.4 to 1.8 mil. Angle of attack was varied from -2 deg to 88 deg at zero sideslip. The purpose of the investigation was to determine the effect of Mach number of the side force that develops at low speeds and zero sideslip for all of these forebody models when the nose is pointed. Test results show that with increasing Mach number the maximum side forces decrease to zero between Mach numbers of 0.8 and 1.5, depending on the nose angle; the smaller the nose angle of the higher the Mach number at which the side force exists. At a Mach number of 0.6 there is some variation of side force with Reynolds number, the variation being the largest for the more slender tangent ogive.

Keener, E. R.↗
NASA NTRS1977

Effect of angle of attack and Mach number on slender wing aerodynamics

An analytic theory is presented in which the classical slender wing theory is modified to account for the combined effects of large angle of attack and nonsonic Mach number on the unsteady aerodynamics. The computed results agree well with available static and dynamic experimental data for slender delta wings in the freestream Mach number range between 0 and 2.8. The method was extended to compute the unsteady aerodynamics of the space shuttle orbiter by defining an equivalent slender wing using static experimental data. The results obtained in this manner are in good agreement with dynamic experimental results for the freestream Mach number range between 0.3 and 1.2.

Ericsson, L. E.↗
NASA NTRS1984

Fuselage and nozzle pressure distributions of a 1/12-scale F-15 propulsion model at transonic speeds. Effect of fuselage modifications and nozzle variables

Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.

Pendergraft, O. C., Jr.↗
NASA NTRS1952

Effects of Aspect Ratio on Air Flow at High Subsonic Mach Numbers

Schlieren photographs were used in an investigation to determine the effects of changing the aspect ratio from infinity to 2 on the air flow past a wing at high subsonic Mach numbers. The results indicated that the decreased effects of compressibility on drag coefficients for the finite wing are produced by a reduction in the compression shock and flow separation.

FLOW, MIXED↗
NASA NTRS1981

Airframe-propulsion system aerodynamic interference predictions at high transonic Mach numbers including off-design engine airflow effects

The transonic speed regime for airplanes at conditions where inlet spillage takes place is discussed. A wind tunnel test program to evaluate aerodynamic performance penalties associated with propulsion system installation and operation at subsonic through low supersonic speeds was conducted. The accuracy of analytic methods for predicting transonic engine airframe interference effects was assessed. Study variables included Mach number, angle of attack, relative nacelle location, and nacelle mass flow ratio. Results include test theory comparisons of forces as well as induced pressure fields. Prediction capability of induced shock wave strength and locations is assessed. It was found that large interference forces due to engine location and flow spillage occur at transonic speeds, that theory explains these effects; and that theory can predict quantitatively these effects.

Kulfan, R. M.↗
NASA NTRS1995

Computational analysis of forebody tangential slot blowing on the high alpha research vehicle

A numerical analysis of forebody tangential slot blowing as a means of generating side force and yawing moment is conducted using an aircraft geometry. The Reynolds-averaged, thin-layer, Navier-Stokes equations are solved using a partially flux-split, approximately-factored algorithm. An algebraic turbulence model is used to determine the turbulent eddy viscosity values. Solutions are obtained using both patched and overset grid systems. In the patched grid model, and actuator plane is used to introduce jet variables into the flow field. The overset grid model is used to model the physical slot geometry and facilitate modeling of the full aircraft configuration. A slot optimization study indicates that a short slot located close to the nose of the aircraft provided the most side force and yawing moment per unit blowing coefficient. Comparison of computed surface pressure with that obtained in full-scale wind tunnel tests produce good agreement, indicating the numerical method and grid system used in the study are valid. Full aircraft computations resolve the changes in vortex burst point due to blowing. A time-accurate full-aircraft solution shows the effect of blowing on the changes in the frequency of the aerodynamic loads over the vertical tails. A study of the effects of freestream Mach number and various jet parameters indicates blowing remains effective through the transonic Mach range. An investigation of the force onset time lag associated with forebody blowing shows the lag to be minimal. The knowledge obtained in this study may be applied to the design of a forebody tangential slot blowing system for use on flight aircraft.

Gee, Ken↗
NASA NTRS2009

Experimental Investigation of the Flow about a 65 deg Delta Wing in the NASA Langley National Transonic Facility

An experimental investigation for the flow about a 65 deg. delta wing has been conducted in the NASA Langley National Transonic Facility (NTF). The tests were conducted at Reynolds numbers, based on the mean aerodynamic chord, ranging from 6 million to 120 million and at Mach numbers ranging from 0.4 to 0.9. The model incorporated four different leading-edge bluntness values. The data include detailed static surfacepressure distributions as well as normal-force and pitching-moment coefficients. The test program was designed to quantify the effects of Mach number, Reynolds number, and leading-edge bluntness on the onset and progression of leading-edge vortex separation.

Luckring, James M.↗