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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.

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At least 1,495 records · Page 83
All sourcesNASA NTRSDOE OSTI
NASA NTRS1956

Lift and Pitching-moment Interference Between a Pointed Cylindrical Body and Triangular Wings of Various Aspect Ratios at Mach Numbers of 1.50 and 2.02

The lift and pitching-moment characteristics of a body alone, six triangular wings of various aspect ratios, and the combinations were measured at Mach numbers of 1.50 and 2.02 at a Reynolds number of 5.5 million (based on the body length) for angles of attack up to 5.5 degrees. The total lift and pitching-moment interference were determined and compared with theory. The agreement was found to be good.

RESEARCH TECHNIQUE - CORRECTIONS↗
NASA NTRS1958

Investigation of Minimum Drag and Maximum Lift-drag Ratios of Several Wing-body Combinations Including a Cambered Triangular Wing at Low Reynolds Numbers and at Supersonic Speeds

Theoretical and experimental data for wing-body combinations with wings of triangular, arrow, and diamond plan form are presented for Mach numbers of 1.62, 1.93, and 2.41 and a Reynolds number range of 1.4 x 10 to the 6th power to 0.7 x 10 to the 6th power. Included are two each of triangular and arrow plan-form wings cambered for approximately uniform load at m = 1.62 and lift coefficients of 0.08 and 0.20. Liquid-film studies of the flow over the various configurations are also presented.

Brown, Clinton E↗
NASA NTRS1949

Investigation of Downwash and Wake Characteristics at a Mach Number of 1.53 I : Rectangular Wing

The results of an experimental investigation of the downwash and wake characteristics behind a rectangular plan-form wing of aspect ratio 3.5 are presented. The airfoil section was a 5-percent-thick, symmetrical double wedge. The tests were made at a Mach number of 1.53 and a Reynolds number of 1.25 million. A comparison between experimental and theoretical values of the downwash angles is made.

STABILITY, LONGITUDINAL - STATIC↗
NASA NTRS1947

A device for measuring sonic velocity and compressor Mach number

A device has been developed which measures the velocity of sound in fluids at stagnation and is especially adaptable to turbine and compressor testing for which the constituency of the working fluid may be in doubt. By utilizing the shaft frequency of a rotary compressor, the instrument can also be used to provide a direct measurement of the compressor Mach number (ratio of blade-tip velocity to inlet velocity of sound at stagnation). A Helmholtz resonator is employed in the measurement of the sound velocity. Viscous effects in the orifice of the Helmholtz resonator are shown to be important and can be taken into account with the help of a parameter obtained from Stokes solution of the flow near an oscillating wall. This parameter includes the kinematic viscosity of the fluid and the frequency of sound in the resonator. When these effects are recognized, the resonator can be calibrated to measure velocity of sound or compressor Mach number to an accuracy of better than 0.5 percent.

Huber, Paul W↗
NASA NTRS1947

An Investigation of Convergent-Divergent Diffusers at Mach Number 1.85

An investigation has been conducted in the Cleveland 18- by 18-inch supersonic tunnel at a Mach number of 1.85 and angles of attack from 0 deg to 5 deg to determine optimum design configurations for a convergent-divergent type of supersonic diffuser with a subsonic diffuser of 5 deg included divergence angle. Total pressure recoveries in excess of theoretical recovery across a normal shock at a free-stream Mach number of 1.85 wore obtained with several configurations. The highest recovery for configurations without a cylindrical throat section was obtained with an inlet having an included convergence angle of 20 deg. Insertion of a 2-inch throat section between a 10 deg included angle inlet and the subsonic diffuser stabilized the shock inside the diffuser and resulted in recoveries as high as 0.838 free-stream total pressure at an angle of attack of 0 deg, corresponding to recovery of 92.4 percent of the kinetic energy of the free air stream. Use of the throat section also lessened the reduction in recovery of all configurations due to angle of attack.

Wyatt, Demarquis D↗
NASA NTRS1950

Aerodynamic Characteristics of NACA RM-10 Missile in 8- by 6-foot Supersonic Wind Tunnel at Mach Numbers from 1.49 to 1.98 II : Presentation and Analysis of Force Measurements

Experimental investigation of aerodynamic forces acting on body of revolution (NACA RM-10 missile) with and without stabilizing fins was conducted at Mach numbers from 1.49 to 1.98 at angles of attack from 0 to 9 degrees and at Reynolds number of approximately 30,000,000. Comparison of experimental lift, drag, and pitching-moment coefficients and center of pressure location for body alone is made with linearized potential theory and a semiempirical method. Results indicate that aerodynamic characteristics were predicted more accurately by semiempirical method than by potential theory. Breakdown of measured drag coefficients into components of friction, pressure, and base-pressure drag is presented for body alone at zero angle of attack.

TAIL-BODY COMBINATIONS - MISSILES↗
NASA NTRS1950

Aerodynamic characteristics of NACA RM-10 missile in 8- by 6-foot supersonic wind tunnel at Mach numbers from 1.49 to 1.98 I : presentation and analysis of pressure measurements (stabilizing fins removed)

Experimental investigation of flow about a slender body of revolution (NACA RM-10 missile) aligned and inclined to a supersonic stream was conducted at Mach numbers from 1.49 to 1.98 at a Reynolds number of approximately 30,000,000. Boundary-layer measurements at zero angle of attack are correlated with subsonic formulations for predicting boundary-layer thickness and profile. Comparison of pressure coefficients predicted by theory with experimental values showed close agreement at zero angle of attack and angle of attack except over the aft leeward side of body. At angle of attack, pitot pressure measurements in plane of model base indicated a pair of symmetrically disposed vortices on leeward side of body.

LOADS - FUSELAGE, NACELLES, AND CANOPIES↗
NASA NTRS1951

Measurements of Downwash and Sidewash Behind Cruciform Triangular Wings at Mach Number 1.4

The results of an experimental investigation of the downwash and sidewash flow angles in a transverse plane behind a cruciform-wing and body combination are presented. The wing-body combination was composed of a cruciform arrangement of two identical triangular wings of aspect ratio 2.31 corresponding to a leading-edge sweep angle of 60°, placed 'on a body of fineness ratio of 16. The investigation was conducted at a Mach number of 1.40 and at a Reynolds number of 1.25 million based on the mean aerodynamic chord of the wings. Dawnwash and sidewash measurements were made for angles of bank of O°, 11-1/4°, 22-1/2°, and 45° and through an angle-of-attack range of ±8O°.

Benton E. Wetzel↗
NASA NTRS1951

Aerodynamic characteristics of four bodies of revolution showing some effects of afterbody shape and fineness ratio at free-stream Mach numbers from 1.50 to 1.99

The effects of fineness ratio (14.2 and 12.2) and boattailing on aerodynamic characteristics of four bodies of revolution at Mach numbers from 1.50 to 1.99 within a range of angles of attack from 0 degrees 10 degrees at an approximate Reynolds number of 35x10(superscript)6 based on body length were investigated. A comparison of experimental data with available theory is included. At zero angle of attack, fineness ratio has no appreciable effect on model characteristics while boattailing and boattail convergence significantly affect fore drag and base drag. At angle of attack the effects are singular. The theory presented by H. J. Allen is a significant improvement over linearized potential theory in predicting aerodynamic characteristics.

BODIES - FITNESS RATIO↗
NASA NTRS1951

Effect of Angle of Attack and Exit Nozzle Design on the Performance of a 16-inch Ram Jet at Mach Numbers from 1.5 to 2.0

An investigation of the performance of a 16-inch ram jet engine having a single oblique-shock all-external compression inlet designed for a flight Mach number of 1.8, was conducted in the NACA Lewis 8-by 6-foot supersonic wind tunnel. Data were obtained at Mach numbers from 1.5 to 2.0 and angles of attack from 0 degrees to 10 degrees. Three exit nozzles were used; a cylindrical extension of the combustion chamber, a 4 degrees half-angle converging nozzle with a 0.71 contraction ratio, and a graphite converging-diverging nozzle having a 0.71 contraction ratio plus reexpansion to essentially major body diameter.

EXITS↗
NASA NTRS1951

Flight Investigation of the Drag of Round-nosed Bodies of Revolution at Mach Numbers from 0.6 to 1.5 Using Rocket-propelled Test Vehicles

Values of total drag coefficient were measured for four round-nosed bodies of revolution in free flight at Mach numbers from 0.6 to 1.5 and Reynolds numbers from 10 x 10(6) to 50 x 10(6). The bodies were designed by rounding off the sharp, fineness-ratio-3.56 nose of a previously tested configuration. The nose radii tested were 27.4, 38.7, 80.6, and 100 percent of the maximum body radius and corresponded to values of 0.075, 0.150, 0.650, and 1.000, respectively, for the ratio of nose-sphere frontal area to body frontal area. The body having least bluntness did not differ appreciably in drag from the pointed-nose body. The others showed marked increases in supersonic drag with increasing bluntness. (author)

Hart, Roger G↗
NASA NTRS1951

Data Presentation of Force Characteristics of Several Engine-strut-body Configurations at Mach Numbers of 1.8 and 2.0 / Robert T. Madden and Emil J. Kremzier

Investigation to determine lift, drag, and pitching-moment characteristics of several engine-strut-body combinations was conducted over range of angles of attack from 0 degrees to 10 degrees at Mach numbers of 1.8 and 2.0. The average Reynolds number based on body length was 28x106. Data are presented without analysis and indicate decreases in minimum drag and lift curve slope with decreasing in minimum drag and lift curve slope with decreasing strut length. Decreases in minimum drag also noted with rear-ward movement of engines.

PROTUBERANCES - BODIES↗
NASA NTRS1951

Flow Separation Ahead of a Blunt Axially Symmetric Body at Mach Numbers 1.76 to 2.10

The pressure distribution and drag were determined for a spherical-nosed axially symmetric body with thin projecting rods at Mach numbers of 1.76, 1.93, and 2.10. The upstream projection distance of the rods was varied over a wide range to study changes in the character of the flow separation and to determine the variation of drag and pressure distribution with tip projection. Drag coefficients between 0.18 and 0.30 were obtained for most tip projections at each Mach number.

WIND TUNNEL TESTS - LFPL - SUPERSONIC (18"x18")↗
NASA NTRS1951

Aerodynamic Characteristics of a Slender Cone-cylinder Body of Revolution at a Mach Number of 3.85

An experimental investigation of the aerodynamics of a slender cone-cylinder body of revolution was conducted at a Mach number of 3.85 for angles of attack of 0 degree to 10 degrees and a Reynolds number of 3.85x10(exp 6). Boundary-layer measurements at zero angle of attack are compared with the compressible-flow formulations for predicting laminar boundary-layer characteristics. Comparison of experimental pressure and force values with theoretical values showed relatively good agreement for small angles of attack. The measured mean skin-friction coefficients agreed well with theoretical values obtained for laminar flow over cones.

FLOW, LAMINAR↗
NASA NTRS1952

Analysis of fully developed turbulent heat transfer at low Peclet numbers in smooth tubes with application to liquid metals

An analysis was made of heat transfer at low Peclet numbers for fluids flowing turbulently in smooth tubes. Previous analyses for flow of gases and liquid metals at low Peclet numbers gave higher heat-transfer coefficients than were indicated experimentally. When the mixing-length theory was modified in order to account for the heat transferred by conduction to a turbulent particle as it moves transversely, the predicted results were brought into agreement with the experimental results.

PIPES↗
NASA NTRS1952

Performance Characteristics of a Normal-shock Side Inlet Located Downstream of a Canard Control Surface at Mach Numbers of 1.5 and 1.8/

The performance characteristics of a downward canted normal-shock side (scoop) inlet located downstream of a triangular control surface are presented for free-stream Mach numbers of 1.5 and 1.8 in terms of total pressure recovery and mass flow ratio for various boundary-layer removal systems,angles of attack, control surface deflections and adverse yaw. An engine operating condition for a hypothetical turbojet engine is established, and the match point characteristics of the engine-inlet configuration are summarized. 520::It is shown that the diffuser performance increases with increased boundary-layer removal and decreases because of the presence of the wake from the forward control surface. At the higher angles of attack the wake passes over the inlet and does not affect the inlet performance. Adverse yaw reduces the total pressure recovery values below those for the unawed case. Magnitudes of the total pressure recovery were below the theoretical normal-shock recovery for the respective test Mach numbers.

INLETS, SIDE - DUCTED BODIES↗
NASA NTRS1952

Spreading of Exhaust Jet from 16 Inch Ream Jet at Mach Number 2.0 / Fred Wilcox, Donald Pennington

An investigation of the jet-spreading characteristics of a 16 inch ram-jet engine was conducted in the 8 by 6 foot supersonic tunnel at a Mach number of 2.0; both a converging nozzle having a contraction ratio of 0.71 and a cylindrical extension to the combustion chamber were used. The jet boundaries determined by means of pitot pressure surveys were compared with boundaries calculated from one-dimensional continuity and momentum relations. For the cylindrical nozzle, the jet reaches its maximum diameter, 4 percent greater than calculated, about 0.6 nozzle-exit diameter downstream of the nozzle exit. The maximum diameter for the converging nozzle was 7 percent greater than calculated from one dimensional relations and occurred from 1 to 1.5 nozzle-exit diameters downstream of the exit. Non dimensional maximum jet diameters agreed closely with results of an investigation by Rousso and Baughman; these data were obtained with low-temperature jets exhausting into a stream at a Mach number of 1.91 from nozzles having exit diameters of 0.75 inch.

INTERFERENCE, JET-MISSILES↗
NASA NTRS1953

Free-flight Performance of a Rocket-boosted, Air-launched 16-inch-diameter Ram-jet Engine at Mach Numbers up to 2.20

The investigation of air-launched ram-jet engines has been extended to include a study of models with a nominal design free-stream Mach number of 2.40. These models require auxiliary thrust in order to attain a flight speed at which the ram jet becomes self-accelerating. A rocket-boosting technique for providing this auxiliary thrust is described and time histories of two rocket-boosted ram-jet flights are presented. In one flight, the model attained a maximum Mach number of 2.20 before a fuel system failure resulted in the destruction of the engine. Performance data for this model are presented in terms of thrust and drag coefficients, diffuser pressure recovery, mass-flow ratio, combustion efficiency, specific fuel consumption, and over-all engine efficiency.

AIR INLETS - NOSE, ANNULAR↗