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  • AERODYNAMICS
  • 1970-1974  (252)
  • 1925-1929
  • 1972  (252)
  • 101
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    A parametric study of planform and aeroelastic effects on aerodynamic center, alpha- and q- stability derivatives. Appendix A: A computer program for calculating alpha- and q- stability derivatives and induced drag for thin elastic aeroplanes at subsonic and supersonic speeds (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The computer program used to determine the rigid and elastic stability derivatives presented in the summary report is listed in this appendix along with instructions for its use, sample input data and answers. This program represents the airplane at subsonic and supersonic speeds as (a) thin surface(s) (without dihedral) composed of discrete panels of constant pressure according to the method of Woodward for the aerodynamic effects and slender beam(s) for the structural effects. Given a set of input data, the computer program calculates an aerodynamic influence coefficient matrix and a structural influence coefficient matrix.
    Keywords: AERODYNAMICS
    Type: NASA-CR-112229 , CRINC-FRL-72-011-APP-A
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  • 102
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    Aerodynamic data obtained from free-flight model firings of two blunt-nosed conical configurations at hypersonic velocities in mixtures of air, nitrogen, and carbon dioxide (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Two blunt-nosed conical configurations were investigated at hypersonic velocities to observe the effect of several gas mixtures on their drag and stability characteristics.
    Keywords: AERODYNAMICS
    Type: NASA-CR-131986 , AD-752764 , NOLTR-72-236
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  • 103
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    A parametric study of planform and aeroelastic effects on aerodynamic center, alpha- and q-stability derivatives. Appendix E: Procedures used to determine the structural representation for idealized low aspect ratio two spar fighter wings (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An explanation is presented of the method used to locate the elastic axis and the method to determine the EI and GJ distributions along the elastic axes of wings with a 2-spar (front and rear) construction or a single torque-box construction.
    Keywords: AERODYNAMICS
    Type: NASA-CR-112233 , CRINC-FRL-72-015-APP-E
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  • 104
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    A parametric study of planform and aeroelastic effects on aerodynamic center, alpha- and q- stability derivatures. Appendix B: Method for computing the strucutral influence coefficient matrix of nonplanar wing body tail configurations (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The method used in computing the structural influence coefficient matrix of the computer program of Reference 1 (appendix A of the Summary Report) is reported. This matrix is computed for complete wing-body-tail configurations by assuming that all major airplane components can be structurally represented by a slender beam called the elastic axis. A structural influence coefficient is defined as the rotation about the Y-stability axis at panel j induced by a unit load on panel k. A description of how a structural breakdown is performed in detail is included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-112230 , CRINC-FRL-72-012-APP-B
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  • 105
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    Effect of nozzle lateral spacing on afterbody drag and performance of twin-jet afterbody models with cone plug nozzles at Mach numbers up to 2.20 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Twin-jet afterbody models were investigated by using two balances to measure separately the thrust minus total drag and the afterbody drag at Mach numbers of 0.0 and 0.50 to 2.20 for a constant angle of attack of 0. Translating shroud cone plug nozzles were tested at dry and maximum afterburning power settings with a high-pressure air system used to provide jet total-pressure ratios up to 20.0. Two nozzle lateral spacings were studied by using afterbodies with several interfairing shapes. The close- and wide-spaced afterbodies had identical cross-sectional area distributions when similar interfairings were installed on each. Nozzle cant angles of -5, 0, and 5 degrees were investigated. The results show that the highest overall performance was generally obtained with the close-spaced afterbody, basic interfairings (no base), and uncanted nozzles.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2632 , L-8481
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  • 106
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    Calculative techniques for transonic flows about certain classes of wing-body combinations, phase 2 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Theoretical analysis and associated computer programs were developed for predicting properties of transonic flows about certain classes of wing-body combinations. The procedures used are based on the transonic equivalence rule and employ either an arbitrarily-specified solution or the local linerization method for determining the nonlifting transonic flow about the equivalent body. The class of wind planform shapes include wings having sweptback trailing edges and finite tip chord. Theoretical results are presented for surface and flow-field pressure distributions for both nonlifting and lifting situations at Mach number one.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2157
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  • 107
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    Flow conditions around the exit and downstream of certain stator blading with various trailing edge thicknesses and geometries (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The blading investigated was curved back with thick profiles. The variations in flow conditions considered were flow angle and isentropic energy. Experimental data were obtained in a two-dimensional cascade from surveys with a combined angle, total-, and static-pressure probe and from an array of end-wall static pressure taps. Analytical results were obtained from ideal flow theory. The results showed large variations in flow conditions close to the plane of the trailing edge that were largely attenuated at a distance a little greater than one blade pitch downstream of the trailing edge in the direction of flow. The results were affected by the geometry and thickness of the trailing edge. The agreement between experimental and analytical results is generally fair to excellent.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2659 , E-7024
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  • 108
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    Heat transfer at reattachment of a compressible flow over a backward facing step with a suction slot. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: A simplified model is suggested for engineering estimates of heat transfer in the vicinity of reattachment for the case of supersonic flow over a backward facing step with a suction slot. The presented approach gives a closed-form approximate solution which is shown to be in satisfactory agreement with experimental results for the case of no mass transfer, and is expected to predict the effect of mass transfer in the case of suction.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 10; Aug. 197
    Format: text
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  • 109
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    An advanced stochastic model for threshold crossing studies of rotor blade vibrations. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: A stochastic model to analyze turbulence-excited rotor blade vibrations, previously described by Gaonkar et al. (1971), is generalized to include nonuniformity of the atmospheric turbulence velocity across the rotor disk in the longitudinal direction. The results of the presented analysis suggest that the nonuniformity of the vertical turbulence over the rotor disk is of little influence on the random blade flapping response, at least as far as longitudinal nonuniformity is concerned.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 10; Aug. 197
    Format: text
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  • 110
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    Response of a trailing vortex to axial injection into the core. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: It is demonstrated that axial air injection into a core of a vortex can beneficially spread out the vorticity concentrated in it and prematurely age it. It is also shown that the phenomenon is more nearly governed by the momentum flux of injection than by mass flow.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 9; Mar. 197
    Format: text
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  • 111
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    Jet effects on the drag of conical afterbodies at supersonic speeds (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A parametric investigation has been conducted to determine the jet effects on the boattail drag of nozzles with truncated conical afterbodies. The boattail drag for nozzle configurations with boattail angles of 3 deg, 5 deg, and 10 deg and ratios of boattail length to maximum diameter of 1.0, 0.8, and 0.6 was compared for the jet-off condition and for a wide range of jet pressure ratios. A nozzle configuration with a circular-arc boattail was tested also. The tests were run at Mach numbers of 1.83 and 2.20 with the model at an angle of attack of 0 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6789 , L-8139
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  • 112
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    Calculation of linearized supersonic flow over slender cones of arbitrary cross section (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Supersonic linearized conical-flow theory is used to determine the flow over slender pointed cones having horizontal and vertical planes of symmetry. The geometry of the cone cross sections and surface velocities are expanded in Fourier series. The symmetry condition permits the uncoupling of lifting and nonlifting solutions. The present method reduces to Ward's theory for flow over a cone of elliptic cross section. Results are also presented for other shapes. Results by this method diverge for cross-sectional shapes where the maximum thickness is large compared with the minimum thickness. However, even for these slender-body shapes, lower order solutions are good approximations to the complete solution.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6818 , L-8192
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  • 113
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    Effects of Leading-Edge Camber on Low-Speed Characteristics of Slender Delta Wings: Techniques and Tabulated Data (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Special tests to determine the importance of transition fixing and Reynolds number variation on forces produced by thin delta wings are discussed. Transition fixing was achieved by applying cement to models and sprinkling grit on wet adhesive. Strips were applied along lines emanating from the apex located along sixty-five percent semi-span rays. Tests were made with grit on both surfaces, upper surface only, lower surface only, and on clean surfaces. Reynolds number varied by testing at three dynamic pressures. Correspondence between dynamic pressure and Reynolds number based on the mean aerodynamic chord are shown. No significant changes are noted due to either Reynolds number variation or transition fixing, within the range of Reynolds numbers used for the test.
    Keywords: AERODYNAMICS
    Type: NASA-CR-112016
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  • 114
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    Experimental investigation of circular, flat, grooved and plain steel diaphragms bursting into a 30.5-centimeter-square section (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Limited data on the bursting of circular, initially flat, grooved and plain steel diaphragms opening into a 30.5-cm-square section are presented in tabular form. In addition, these data were used to determine values of an empirical constant to be used in a design equation for predicting diaphragm bursting pressures and opening times.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2549 , L-4645
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  • 115
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    A method for analyzing dynamic stall of helicopter rotor blades (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A model for each of the basic flow elements involved in the unsteady stall of a two-dimensional airfoil in incompressible flow is presented. The interaction of these elements is analyzed using a digital computer. Computations of the loading during transient and sinusoidal pitching motions are in good qualitative agreement with measured loads. The method was used to confirm that large torsional response of helicopter blades detected in flight tests can be attributed to dynamic stall.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2009
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  • 116
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    Acoustic testing of a 1.5 pressure ratio, low tip speed fan (QEP fan B scale model) (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A scale model (0.484 scale factor) of a single stage fan designed for a 1.5 pressure ratio and 1160 ft/sec tip speed was tested to determine its noise characteristics. The fan had 26 blades and 60 outlet guide vanes, with vanes spaced two rotor blade aerodynamic chords from the blades. The effects of speed, exhaust nozzle area and fan frame acoustic treatment on the scale model's noise characteristics were investigated.
    Keywords: AERODYNAMICS
    Type: NASA-CR-120789
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  • 117
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    Lifting-surface theory for calculating the loading induced on a wing by a flap (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A method is described for using lifting-surface theory to obtain the pressure distribution on a wing with a trailing-edge flap or control surface. The loading has a logarithmic singularity at the flap edges, which may be determined directly by the method of matched asymptotic expansions. Expressions are given for the singular flap loading for various flap hinge line and side edge geometries, both for steady and unsteady flap deflection. The regular part of the flap loading must be obtained by inverting the lifting-surface-theory integral equation relating the pressure and the downwash on the wing: procedures are described to accomplish this for a general wing and flap geometry. The method is applied to several example wings, and the results are compared with experimental data. Theory and test correlate well.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6798 , A-4084
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  • 118
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    Inviscid analysis of jet injection between two moving streams (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An analytical method is developed for determining the flow interaction when a two-dimensional jet is injected between two moving streams. The jet is flowing out of channel and is turned as it enters between the external streams. The local velocity variation resulting from the flow interaction provides a static pressure variation along the jet bounding streamlines that is a priori unknown. Hense, the flow must be obtained by coupling the three flow regions (the jet and the free stream on either side) along the jet boundaries. Both external streams have the same total pressure, which is different from that in the jet. The solution is for the condition that the total pressure in the jet does not differ from the free-stream value by a large amount compared with the free-stream dynamic head. Results are given for the shape of the jet boundaries for various injection configurations.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6780 , E-6583
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  • 119
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    Exploratory investigation of lift induced on a swept wing by a two-dimensional partial-span deflected jet at Mach numbers from 0.20 to 1.30 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An exploratory investigation was conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.20 to 1.30 to determine the induced lift characteristics of a body and swept-wing configuration having a partial-span two-dimensional propulsive nozzle with exhaust exit in the notch of the swept-wing trailing edge. The Reynolds number per meter varied from 4,900,000 to 14,030,000. The effects on wing-body characteristics of deflecting the propulsive jet in the flap mode at nominal exhaust-nozzle deflection angles of 0 deg and 30 deg were studied for two nozzle designs with different geometry and wing spans.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2529 , L-8177
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  • 120
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    Exploratory investigation of factors affecting the wing tip vortex (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted in the Langley full-scale tunnel to study some factors affecting the tip vortex of a wing. It was found that there was a pronounced effect of Reynolds number on the tip-vortex core size. An attempt was made to determine what aerodynamic parameters, such as lift, drag, or induced drag, influence the size of the vortex core, but no particular function of the parameters was found to be superior to all others. Various spoilers placed on the upper and lower surfaces of the wing to increase the boundary-layer thickness resulted in a reduction in the vorticity as determined from the tuft grid. Various solid objects placed in the vortex core downstream of the wing tip seemed to decrease the vorticity within the vortex core.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2516 , L-8104
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  • 121
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    Analytical correlation of centrifugal compressor design geometry for maximum efficiency with specific speed (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Centrifugal compressor performance was examined analytically to determine optimum geometry for various applications as characterized by specific speed. Seven specific losses were calculated for various combinations of inlet tip-exit diameter ratio, inlet hub-tip diameter ratio, blade exit backsweep, and inlet-tip absolute tangential velocity for solid body prewhirl. The losses considered were inlet guide vane loss, blade loading loss, skin friction loss, recirculation loss, disk friction loss, vaneless diffuser loss, and vaned diffuser loss. Maximum total efficiencies ranged from 0.497 to 0.868 for a specific speed range of 0.257 to 1.346. Curves of rotor exit absolute flow angle, inlet tip-exit diameter ratio, inlet hub-tip diameter ratio, head coefficient and blade exit backsweep are presented over a range of specific speeds for various inducer tip speeds to permit rapid selection of optimum compressor size and shape for a variety of applications.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6729 , E-6638
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  • 122
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    A wind-tunnel evaluation of analytical techniques for predicting static stability and control characteristics of flexible aircraft (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental evaluation of analytical techniques for predicting certain stability and control characteristics of a large flexible aircraft is presented. Analytical methods based on both the model approach and flexibility influence coefficients are developed to predict the aerodynamic characteristics of a flexible airplane. These methods are then applied to a flexibly scaled model of a supersonic transport configuration. Comparisons of wind-tunnel data, calculations based on the model approach, and flexibility influence coefficients are presented over the Mach number range from 0.6 to 2.7. An examination of the results obtained from this study indicates that both analytical techniques predict reasonably well the effect of flexibility on the basic longitudinal characteristics and that both techniques give generally comparable results.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6656 , L-8105
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  • 123
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    Performance of 1380-foot-per-second tip-speed axial-flow compressor rotor with blade tip solidity of 1.1 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The aerodynamic design parameters are presented along with the overall and blade element performance, of an axial-flow compressor rotor designed to study the effects of blade solidity on efficiency and stall margin. At design speed the peak efficiency was 0.853 and occured at an equivalent weight flow of 65.7lb/sec. The total pressure ratio was 1.68. Design efficiency, weight flow, pressure ratio, and temperature ratio were 0.822, 65.3, 1.65, and 1.187, respectively. Stall margin for design speed was 14 percent based on the weight flows and pressure ratios at peak efficiency and just prior to stall.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2449 , E-6686
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  • 124
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    Wind-tunnel simulation of store jettison with the aid of magnetic artificial gravity (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A method employed in the simulation of jettison of stores from aircraft involving small scale wind-tunnel drop tests from a model of the parent aircraft is described. Proper scaling of such experiments generally dictates that the gravitational acceleration should ideally be a test variable. A method of introducing a controllable artificial component of gravity by magnetic means has been proposed. The use of a magnetic artificial gravity facility based upon this idea, in conjunction with small scale wind-tunnel drop tests, would improve the accuracy of simulation. A review of the scaling laws as they apply to the design of such a facility is presented. The design constraints involved in the integration of such a facility with a wind tunnel are defined. A detailed performance analysis procedure applicable to such a facility is developed. A practical magnet configuration is defined which is capable of controlling the strength and orientation of the magnetic artificial gravity field in the vertical plane, thereby allowing simulation of store jettison from a diving or climbing aircraft. The factors involved in the choice between continuous or intermittent operation of the facility, and the use of normal or superconducting magnets, are defined.
    Keywords: AERODYNAMICS
    Type: NASA-CR-1955 , TR-174
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  • 125
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    Flight-determined characteristics of an air intake system on an F-111A airplane (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Flow phenomena of the F-111A air intake system were investigated over a large range of Mach number, altitude, and angle of attack. Boundary-layer variations are shown for the fuselage splitter plate and inlet entrance stations. Inlet performance is shown in terms of pressure recovery, airflow, mass-flow ratio, turbulence factor, distortion factor, and power spectral density. The fuselage boundary layer was found to be not completely removed from the upper portion of the splitter plate at all Mach numbers investigated. Inlet boundary-layer ingestion started at approximately Mach 1.6 near the translating spike and cone. Pressure-recovery distribution at the compressor face showed increasing distortion with increasing angle of attack and increasing Mach number. The time-averaged distortion-factor value approached 1300, which is near the distortion tolerance of the engine at Mach numbers above 2.1.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6679 , H-661
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  • 126
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    Effect of limited amplitude and rate of flap motion on vane-controlled gust alleviation system (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An airplane (light transport type) is assumed to be in level flight (no pitching) through atmospheric turbulence which has a mean-square vertical gust intensity of 9.3 (m/sec)sq. The power spectral density of the vertical acceleration due to gusts is examined with and without a gust-alleviation system in operation. The gust-alleviation system consisted of wing flaps that were used in conjunction with a vane mounted ahead of the airplane to sense the vertical gust velocity. The primary purpose of this study was to examine the change in the effectiveness of the gust-alleviation system when the flap motion is limited in amplitude and rate. The alleviation system was very effective if no restrictions were placed on flap motion (rate and amplitude). Restricting the flap amplitude to 0.5 radian did not appreciably change the effectiveness. However, restricting the flap rate did reduce the gust alleviation, and restricting the flap rate to 0.25 rad/sec actually caused the alleviation system to increase the vertical acceleration above that for the no-alleviation situation. Based upon this analysis, rate limiting appears to be rather significant in gust-alleviation systems designed for passenger comfort.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6733 , L-8027
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  • 127
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    Design and experimental results of a highly loaded, low solidity, jet flap rotor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The overall performance of a single-stage turbine with a low solidity jet flap rotor blade assembly was tested over a range of cavity pressure ratios, equivalent speeds, and expansion ratios. The rotor blades were designed with negative hub reaction and a mean-line axial chord solidity of 0.922. The results of the investigation are compared with the performance of a modified jet flap rotor blade which was designed to similar velocity diagrams but with a mean-section, axial chord solidity of 1.541. Both rotors were tested with the same stator.
    Keywords: AERODYNAMICS
    Type: NASA-CR-1968 , EDR-7045
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  • 128
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    Aerodynamic characteristics of an NASA supercritical-wing research airplane model with and without fuselage area-rule additions at Mach 0.25 to 1.00 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Transonic pressure tunnel tests at Mach numbers from 0.25 to 1.00 were performed to determine the effects of area-rule additions to the sides of the fuselage on the aerodynamic characteristics of a 0.087 scale model of an NASA supercritical-wing research airplane. Presented are the longitudinal aerodynamic force and moment characteristics for horizontal-tail deflection angles of -2.5 deg and -5 deg with the side fuselage area-rule additions on and off the model. The effects of the side fuselage area-rule additions on selected wing and fuselage pressure distributions at near-cruise conditions are also presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2633 , L-8422
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  • 129
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    Wind-tunnel measurements of aerodynamic load distribution on an NASA supercritical-wing research airplane configuration (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Wind tunnel tests have been conducted on a research airplane model with an NASA supercritical wing to define the general character of the flow over the wing and to aid in structural design of the full scale airplane. Pressure measurements were made at Mach numbers from 0.25 to 1.30 for sideslip angles from -2.50 deg to 2.50 deg over a moderate range of angles of attack and dynamic pressures. Except for representative figures, the results are presented in tabular form without detailed analysis.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2469 , L-7982
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  • 130
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    Wind-tunnel investigation of basic aerodynamic characteristics of a supercritical-wing research airplane configuration (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Transonic pressure tunnel and transonic tunnel tests were performed to determine the aerodynamic characteristics of a 0.087 scale model of a supercritical wing research airplane configuration at Mach numbers from 0.25 to 1.30. The investigation included tests to determine the basic longitudinal aerodynamic characteristics, the lateral-directional aerodynamic characteristics for sideslip angles of 0 deg and + or - 2.5 deg, and the effects of Reynolds number and aeroelasticity.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2470 , L-7979
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  • 131
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    Tabulated pressure measurements of a NASA supercritical-wing research airplane model with and without fuselage area-rule additions at Mach 0.25 to 1.00 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Basic pressure measurements were made on a 0.087-scale model of a supercritical wing research airplane in the Langley 8 foot transonic pressure tunnel at Mach numbers from 0.25 to 1.00 to determine the effects on the local aerodynamic loads over the wing and rear fuselage of area-rule additions to the sides of the fuselage. In addition, pressure measurements over the surface of the area-rule additions themselves were obtained at angles of sideslip of approximately - 5 deg, 0 deg, and 5 deg to aid in the structural design of the additions. Except for representative figures, results are presented in tabular form without analysis.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2634 , L-8443
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  • 132
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    Feasibility study of inlet shock stability system of YF-12 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The feasibility of self actuating bleed valves as a shock stabilization system in the inlet of the YF-12 is considered for vortex valves, slide valves, and poppet valves. Analytical estimation of valve performance indicates that only the slide and poppet valves located in the inlet cowl can meet the desired steady state stabilizing flows, and of the two the poppet valve is substantially faster in response to dynamic disturbances. The poppet valve is, therefore, selected as the best shock stability system for the YF-12 inlet.
    Keywords: AERODYNAMICS
    Type: NASA-CR-134594 , SP-1964
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  • 133
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    Unsteady rotor aerodynamics at low inflow and its effect on flutter. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: A comprehensive flutter analysis has been developed utilizing compressible unsteady aerodynamics for a helicopter rotor at a low inflow condition. The rotor is treated as a nonuniform rotating beam with finite bending and torsional stiffness. The unsteady aerodynamic representation incorporates the effects of spanwise variations in Mach number and reduced frequency by applying available two-dimensional theories in a strip theory fashion. The results are characterized by an oscillation of the apparent aerodynamic damping with decreasing frequency ratio (oscillation frequency/rotational frequency) which suggests a new flutter criteria for rotors. These results are correlated with available experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-959 , American Institute of Aeronautics and Astronautics, Atmospheric Flight Mechanics Conference; Sept. 11-13, 1972; Palo Alto, CA
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  • 134
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    Installation caused flow distortion and its effect on noise from a fan designed for turbofan engines. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: Use of a ground test stand to obtain acoustic data on a full-scale prototype fan designed for quiet subsonic-aircraft engines. The fan was installed in three different ways in the test stand. In two of the installations the fan was driven by a shaft in the inlet; in the third installation the fan was driven from the rear. These three installations, and the structures associated with them, resulted in various amounts of inlet flow distortion to the fan. The rear-drive installation had less inlet flow distortion than the two front drive installations. Differences in blade passage sound pressure levels of more than 10 dB were measured between the rear-drive and front-drive versions, with the rear-drive installation producing less noise. Perceived noise levels were computed and the influence of the distortion on these levels was determined.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-1006 , American Institute of Aeronautics and Astronautics, Aerodynamic Testing Conference; Sept. 13-15, 1972; Palo Alto, CA
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  • 135
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    Prediction of far flow field in trailing vortices. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: A finite-difference machine code is brought to bear on the wake-vortex problem in the quasi-cylindrical boundary-layer approximation. A turbulent-energy model containing new features is developed. Parameters of the model are evaluated by comparison of calculated velocities and turbulent intensities with measurements in an axisymmetric wake. Comparisons are made with a previous calculation of the decay of an isolated vortex and with wind tunnel and flight measurements in trailing vortices. A self-similar solution develops at large axial distance that decays with the square root of distance. A slower decay occurs in the preceding transition region.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-989 , American Institute of Aeronautics and Astronautics, Atmospheric Flight Mechanics Conference; Sept. 11-13, 1972; Palo Alto, CA
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  • 136
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    Aerodynamics of high-performance Ludwieg tubes. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: A small perturbation analysis of the flow in a Ludwieg tube supply is described which includes both the growing turbulent boundary-layer and axial pipe nonuniformities. Measured pressure time histories are used to check the theory and its extension to high pipe Mach numbers and large flow perturbations. The nonsteady coupling of this flow to a nozzle is investigated. Calculations are made of the effects on the nozzle boundary layer, while a separate analysis is carried out for the case of complete mixing in a nozzle plenum. The double expansion nozzle is shown to extend Ludwieg-tube capability. Analytical and experimental studies are made of the aerodynamic problems that can arise in these nozzles.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-994 , American Institute of Aeronautics and Astronautics, Aerodynamic Testing Conference; Sept. 13-15, 1972; Palo Alto, CA
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  • 137
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    Initial rotation-loading and low speed flutter test results for a straight wing version of the space shuttle vehicle (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For three straight semispan model space shuttle wings, the maximum total load during rapid rotation from 66 deg to 0 deg angle of attack, at Mach numbers from 0.28 to 0.60, was essentially no higher than that measured for buffet. During slow rotation over the same angle range, there was no visible flutter. For one of the wings, however, unstable aerodynamic damping was established at two fixed angles of attack.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62110
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  • 138
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    Design and performance of an 0.8 hub-tip ratio axial flow pump rotor with a blade tip diffusion factor of 0.55 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 22.9-centimeter diameter axial flow rotor with a 0.8 hub-tip radius ratio, a design flow coefficient of 0.466, and a blade tip design diffusion factor of 0.55 was tested in cold water under both cavitating and noncavitating conditions. Radial surveys of the flow conditions at the rotor inlet and outlet were made. At design flow, the rotor produced an overall headrise coefficient of 0.360 with an overall efficiency of 95.0 percent. The efficiency remained greater than 88 percent over the entire flow coefficient range which varied from 0.350 to 0.615.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2485 , E-6423
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  • 139
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    Subsonic and supersonic static aerodynamic characteristics of a family of bulbous base cones measured with a magnetic suspension and balance system (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of subsonic and supersonic wind-tunnel tests with a magnetic balance and suspension system on a family of bulbous based cone configurations are presented. At subsonic speeds the base flow and separation characteristics of these configurations is shown to have a pronounced effect on the static data. Results obtained with the presence of a dummy sting are compared with support interference free data. Support interference is shown to have a substantial effect on the measured aerodynamic coefficient.
    Keywords: AERODYNAMICS
    Type: NASA-CR-1932 , TR-166
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  • 140
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    Sideslip of wing-body combinations (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A small-disturbance theory is developed for predicting the aerodynamics of an airplane in sideslip. Second-order terms involving the interaction between sideslip angle and angle of attack, sideslip angle and wing camber, etc., are retained. It is found that the second-order terms can produce the dominant sideslip effects when the dihedral of the lifting surfaces is small. Numerical implementation of the theory requires a solution procedure capable of producing accurate velocity gradients in the first-order solution.
    Keywords: AERODYNAMICS
    Type: NASA-CR-114716 , D6-60160
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  • 141
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    Investigation of a mixed compression axisymmetric inlet at Mach number 5.3 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The hypersonic diffuser portion of an uncooled high performance mixed compression, axisymmetric inlet suitable for subsonic burning engines was designed and tested. Performance of a model with a 25.4-cm capture diameter was measured in a wind tunnel and the results were compared with theoretical predictions calculated by a comprehensive computer program. All tests were conducted at a Mach number of 5.3 at a total temperature of 667 K and a total pressure of 11.57 atm. The angle of attack ranged from 0 to + or - 3 deg. Performance at angle of attack remained high. Reasonably high performance in the throat (maximum throat pitot-pressure recovery of 77 percent and an average value of 58 percent) was obtained at 0 deg angle of attack with relatively large amounts of boundary-layer bleed (11 to 22 percent of the capture mass flow). The computer program used in the design of this inlet is considered marginally adequate for predicting hypersonic inlet flow fields. Although the program as it now exists is very useful, an improved computer program that more accurately predicts the boundary layer and the shock-wave-boundary-layer interaction and accounts for boundary-layer bleed should be developed for reliability predicting hypersonic inlet flow fields.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6647 , A-4160
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  • 142
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    Reynolds number effect on overall performance of a 10.8-centimeter (4.25-inch) sweptback bladed centrifugal compressor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The performance data were taken at 50,000 rpm, using argon gas. As the Reynolds number was reduced from near design value to 30 percent of design, the maximum efficiency decreased about 1.5 percentage points. Reducing the Reynolds number from 30 percent to approximately 10 percent of design caused the maximum efficiency to decrease another 2.5 percentage points. The variation in loss with Reynolds number is compared with inverse power relation of loss with Reynolds number.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6640 , E-6586
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  • 143
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    Wind tunnel study of slot spoilers for direct lift control (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted in a 300-mph 7- by 10- foot tunnel to obtain data for a slot spoiler direct lift control system. Slot spoilers are believed to have advantages over flap-type direct lift control (DLC) systems because of the small amount of power required for actuation. These tests, run at a Reynolds number of 1,400,000 showed that up to 78 percent of the lift due to flap deflection could be spoiled by opening several spanwise slots within the flaps. For a given lift change the drag change was significantly less than that which would be obtained by a variable flap DLC system. A nozzle-shaped slot was the most effective of the slot shapes tested.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6627 , L-8014
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  • 144
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    Effect of Reynolds number on overall performance of a 3.7-inch-diameter six-stage axial-flow compressor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 9.4-centimeter (3.7-in.) diameter six-stage axial-flow compressor was tested in argon over a range of inlet pressures corresponding to a Reynolds number range of 30,600 to 160,000. The effect of Reynolds number on efficiency, pressure ratio, work input, maximum flow, and surge is shown. The Reynolds number effects are discussed in terms of changes in boundary-layer thickness, losses, and the resulting changes in throughflow velocity. Significant deviation was noted from the 0.2 power relation often used to express the variation of loss with Reynolds number.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6628 , E-6522
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  • 145
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    Surface pressure and inviscid flow field properties of the McDonnell-Douglas delta-wing orbiter for nominal Mach number of 8, Volume 1 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of a wind tunnel test program to determine the surface pressures and flow distribution on the McDonnell Douglas Orbiter configuration are presented. Tests were conducted in hypersonic wind tunnel at Mach 8. The freestream unit Reynolds number was 3.7 time one million per foot. Angle of attack was varied from 10 degrees to 60 degrees in 10 degree increments.
    Keywords: AERODYNAMICS
    Type: NASA-CR-120037-VOL-1 , DMS-DR-1225-VOL-1 , DRC-184-58
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  • 146
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    Exploratory investigation of sound pressure level in the wake of an oscillating airfoil in the vicinity of stall (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Wind tunnel tests were performed on two oscillating two-dimensional lifting surfaces. The first of these models had an NACA 0012 airfoil section while the second simulated the classical flat plate. Both of these models had a mean angle of attack of 12 degrees while being oscillated in pitch about their midchord with a double amplitude of 6 degrees. Wake surveys of sound pressure level were made over a frequency range from 16 to 32 Hz and at various free stream velocities up to 100 ft/sec. The sound pressure level spectrum indicated significant peaks in sound intensity at the oscillation frequency and its first harmonic near the wake of both models. From a comparison of these data with that of a sound level meter, it is concluded that most of the sound intensity is contained within these peaks and no appreciable peaks occur at higher harmonics. It is concluded that within the wake the sound intensity is largely pseudosound while at one chord length outside the wake, it is largely true vortex sound. For both the airfoil and flat plate the peaks appear to be more strongly dependent upon the airspeed than on the oscillation frequency. Therefore reduced frequency does not appear to be a significant parameter in the generation of wake sound intensity.
    Keywords: AERODYNAMICS
    Type: NASA-CR-1948
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  • 147
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    Application of a one-strip integral method to the unsteady supersonic aerodynamics of an inclined flat surface (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The method of integral relations is applied in a one-strip approximation to the perturbation equations governing small motions of an inclined, sharp-edged, flat surface about the mean supersonic steady flow. Algebraic expressions for low reduced-frequency aerodynamics are obtained and a set of ordinary differential equations are obtained for general oscillatory motion. Results are presented for low reduced-frequency aerodynamics and for the variation of the unsteady forces with frequency. The method gives accurate results for the aerodynamic forces at low reduced frequency which are in good agreement with available experimental data. However, for cases in which the aerodynamic forces vary rapidly with frequency, the results are qualitatively correct, but of limited accuracy. Calculations indicate that for a range of inclination angles near shock detachment such that the flow in the shock layer is low supersonic, the aerodynamic forces vary rapidly both with inclination angle and with reduced frequency.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6644 , L-8047
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  • 148
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    Comparison of hinge moments for a simple delta wing and a delta-wing orbiter concept at Mach 6 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Comparisons of hinge moments for simple delta wing and delta wing orbiter concept at Mach 6
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6657 , L-8103
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  • 149
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    Main rotor free wake geometry effects on blade air loads and response for helicopters in steady maneuvers. Volume 2: Program listings (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A mathematical model and computer program was implemented to study the main rotor free wake geometry effects on helicopter rotor blade air loads and response in steady maneuvers. Volume 1 (NASA CR-2110) contains the theoretical formulation and analysis of results. Volume 2 contains the computer program listing.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2111 , RASA-71-13-VOL-2
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  • 150
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    Performance of a single-stage transonic compressor with a blade-tip solidity of 1.3 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design and experimental performance of a 50-centimeter-diameter, single stage, axial flow, transonic compressor with a blade tip solidity of 1.3 are presented. Radial surveys were made of the flow conditions for both the rotor and stator. At design speed, peak efficiencies for both rotor and stage were 0.87 and 0.82, respectively, and occurred at an equivalent weight flow of 29.6 kilograms per second (202 kg/sec/sq m of annulus area). At peak efficiency, the total pressure ratios for both rotor and stage were 1.79 and 1.73, respectively. Stall margin for the stage was 17 percent.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2645 , E-6763
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  • 151
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    Main rotor free wake geometry effects on blade air loads and response for helicopters in steady maneuvers. Volume 1: Theoretical formulation and analysis of results (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A mathematical model and computer program were implemented to study the main rotor free wake geometry effects on helicopter rotor blade air loads and response in steady maneuvers. The theoretical formulation and analysis of results are presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2110 , RASA-71-13-VOL-1
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  • 152
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    Comparison of aerogrids and punched plates for smoothing flow from short annular diffusers (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Scale model tests were conducted to evaluate the effectiveness of aerogrids and punched plates in producing flat velocity profiles downstream of short diffusers as would be used between the compressor and combustor of advanced aircraft engines. The diffuser had an area ratio of 4.17 and a length-to-inlet-height ratio of 2.07. The aerogrids tested were plates containing 1123 contoured venturis in parallel with geometric blockages of 83, 74, and 61 percent, respectively. The punched plates contained 1123 sharp-edged orifices with blockages of 58 and 30 percent. The results show that aerogrids, with higher effective blockage for the same pressure loss, are more effective flow-smoothing devices than the punched plates. Also, the overall pressure loss decreases and the exit velocity profile becomes flatter as either type of grid is moved closer to the diffuser exit plane.
    Keywords: AERODYNAMICS
    Type: NASA-CR-120960
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  • 153
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    Drag characteristics of a disk-gap-band parachute with a nominal diameter of 1.65 meters at Mach numbers from 2.0 to 3.0 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Supersonic wind-tunnel tests were conducted with disk-gap-band parachute models having a nominal diameter of 1.65 meters and geometric porosities of 10.0, 12.5, and 15.0 percent. Canopy inflation characteristics, angles of attack, and drag performance are presented for deployment behind forebody base extensions which were free to oscillate in pitch and yaw. The effect of increasing suspension-line length on canopy motions and drag performance is included, and the drag performance of a model with 12.5 percent geometric porosity is compared with results from flight tests of a parachute with a nominal diameter of 12.19 meters.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6894 , L-8357
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  • 154
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    An efficient algorithm using matrix methods to solve wind tunnel force-balance equations (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An iterative procedure applying matrix methods to accomplish an efficient algorithm for automatic computer reduction of wind-tunnel force-balance data has been developed. Balance equations are expressed in a matrix form that is convenient for storing balance sensitivities and interaction coefficient values for online or offline batch data reduction. The convergence of the iterative values to a unique solution of this system of equations is investigated, and it is shown that for balances which satisfy the criteria discussed, this type of solution does occur. Methods for making sensitivity adjustments and initial load effect considerations in wind-tunnel applications are also discussed, and the logic for determining the convergence accuracy limits for the iterative solution is given. This more efficient data reduction program is compared with the technique presently in use at the NASA Langley Research Center, and computational times on the order of one-third or less are demonstrated by use of this new program.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6860 , L-8278
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  • 155
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    Investigation of the Applicability of the Free-wing Principle to Light, General Aviation Aircraft (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Gust-alleviation benefits for aircraft employing an unconventional wing, free to pivot about a spanwise axis forward of its aerodynamic center and subject only to aerodynamic pitching moments imposed by lift and drag forces and a trailing-edge control surface are reviewed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2046
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  • 156
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    Summary of transformation equations and equations of motion used in free flight and wind tunnel data reduction and analysis (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Basic formulations for developing coordinate transformations and motion equations used with free-flight and wind-tunnel data reduction are presented. The general forms presented include axes transformations that enable transfer back and forth between any of the five axes systems that are encountered in aerodynamic analysis. Equations of motion are presented that enable calculation of motions anywhere in the vicinity of the earth. A bibliography of publications on methods of analyzing flight data is included.
    Keywords: AERODYNAMICS
    Type: NASA-SP-3070
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  • 157
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    Experimental static aerodynamics of a regular hexagonal prism in a low density hypervelocity flow (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A regular hexagonal prism, having a fineness ratio of 1.67, has been tested in a wind tunnel to determine its static aerodynamic characteristics in a low-density hypervelocity flow. The prism tested was a 1/4-scale model of the graphite heat shield which houses the radioactive fuel for the Viking spacecraft auxiliary power supply. The basic hexagonal prism was also modified to simulate a prism on which ablation of one of the six side flats had occurred. This modified hexagonal prism was tested to determine the effects on the aerodynamic characteristics of a shape change caused by ablation during a possible side-on stable reentry.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6816 , L-8229
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  • 158
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    Aerodynamic characteristics of a 60 degree swept delta-wing space shuttle orbiter at Mach numbers of 2.50, 3.90, and 4.60 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Wind tunnel tests were performed to determine the static aerodynamic characteristics of a model of a 60 degree swept delta wing space shuttle orbiter. Some control effectiveness tests were included. The tests were conducted in a unitary plan wind tunnel at Mach numbers of 2.50, 3.90, and 4.60.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2561 , L-8243
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  • 159
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    Aeronautical Engineering: A special bibliography with indexes, supplement 16 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: This special bibliography lists 426 reports, articles, and other documents introduced into the NASA scientific and technical information system in February 1972.
    Keywords: AERODYNAMICS
    Type: NASA-SP-7037(16)
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  • 160
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    Supersonic aerodynamic characteristics of a low-aspect-ratio missile model with wing and tail controls and with tails in line and interdigitated (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A study has been made to determine the aerodynamic characteristics of a low-aspect ratio cruciform missile model with all-movable wings and tails. The configuration was tested at Mach numbers from 1.50 to 4.63 with the wings in the vertical and horizontal planes and with the wings in a 45 deg roll plane with tails in line and interdigitated.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2531 , L-8092
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  • 161
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    FORTRAN program for generating a two-dimensional orthogonal mesh between two arbitrary boundaries (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A FORTRAN 4 program is presented which computes and plots coordinates for a two-dimensional orthogonal mesh in the region between the walls of a flow channel. The program is designed for a channel containing a body about which flow passes and which spans the channel from one wall to the other. However, the condition that the channel contain an immersed body can be easily removed from the program. Input to the program consists of spline points of the channel walls and the body geometry. Output includes printed and plotted coordinates of the generated orthogonal mesh and angles of the mesh with the horizontal plane.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6766 , E-6644
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  • 162
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    Multiple-orifice liquid injection into hypersonic airstreams and applications to ram C-3 flight (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Experimental data are presented for the oblique injection of water and three electrophilic liquids (fluorocarbon compounds) through multiple-orifice nozzles from a flat plate and the sides of a hemisphere-cone (0.375 scale of RAM C spacecraft) into hypersonic airstreams. The nozzle patterns included single and multiple orifices, single rows of nozzles, and duplicates of the RAM C-III nozzles. The flat-plate tests were made at Mach 8. Total pressure was varied from 3.45 MN/m2 to 10.34 MN/m2, Reynolds number was varied form 9,840,000 per meter to 19,700,000 per meter, and liquid injection pressure was varied from 0.69 MN/m2 to 3.5 MN/m2. The hemisphere-cone tests were made at Mach 7.3. Total pressure was varied from 1.38 MN/m2, to 6.89 MN/m2, Reynolds number was varied from 3,540,000 per meter to 17,700,000 per meter, and liquid-injection pressure was varied from 0.34 MN/m2 to 4.14 MN/m2. Photographs of the tests and plots of liquid-penetration and spray cross-section area are presented. Maximum penetration was found to vary as the square root of the dynamic-pressure ratio and the square root of the total injection nozzle area. Spray cross-section area was linear with maximum penetration. The test results are used to compute injection parameters for the RAM C-3 flight injection experiment.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2486 , L-8023
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  • 163
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    Performance of 1380 foot per second tip-speed axial-flow compressor rotor blade tip solidity of 1.5 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: This presents the aerodynamic design parameters along with the overall and blade element performance of an axial-flow compressor rotor designed to study the effects of blade solidity on efficiency and stall margin. At design speed the peak efficiency was 0.892 and occurred at an equivalent weight flow of 65.0 lb/sec. The total pressure ratio was 1.83 and the total temperature ratio was 1.215. Design efficiency, weight flow, pressure ratio, and temperature ratio were 0.824, 65.3, 1.65, and 1.187, respectively. Stall margin for design speed was 10 percent based on the weight flow and pressure ratio values at peak efficiency and just prior to stall.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2379 , E-5723
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  • 164
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    Effect of trailing edge geometry and thickness on the performance of certain turbine stator blading (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The experimental and analytical investigation included solid blades with five different trailing-edge thicknesses and four different trailing-edge geometries. One of the geometries was round, one was square, one was tapered from the suction surface, and the other tapered from the pressure surface. One of the trailing-edge thicknesses was sharp edged; the other four thicknesses were equivalent to about 5, 11, 16, and 20 percent of the blade throat width. The experimental results show increased efficiency loss for increased trailing-edge thickness for all trailing-edge geometries. The blade with round trailing edge, equal to about 11 percent of the blade throat width, had 60 percent more loss than the sharp-edged blade. For the same trailing-edge thickness, square trailing edges caused more loss than round trailing edges, and the tapered trailing edges caused about the same loss as the round trailing edges.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6637 , E-6613
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  • 165
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    Effect of grooved casing treatment on the flow range capability of a single-stage axial-flow compressor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Comparisons of the results of testing a single-stage axial-flow compressor with a solid-wall casing and with grooved casings are presented. The depth, location, and number of circumferential grooves in the casing over the casing over the rotor tip were varied. The near-stall weight flow was lower than that with the solid-wall casing for all but one grooved configuration indicating an improvement in the stall margin. The greatest reduction in the near-stall weight flow was noted for the configuration with five grooves located over the blade midchord region.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2459 , E-6560
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  • 166
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    Effects of Reynolds number and body corner radius on aerodynamic characteristics of a space shuttle-type vehicle at subsonic Mach numbers (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Static aerodynamic forces and moments were measured to study the effects of Reynolds number and body corner radius on the aerodynamic characteristics of a straight wing space shuttle orbiter at subsonic speeds. A 0.02-scale model was tested at Mach numbers from 0.3 to 0.9 and Reynolds numbers from about 600,000 to 3 million, based on body width. The body alone and the body with its wing and horizontal tail attached were tested at angles of attack from 35 to 75 degrees. The effects of rounding the body corners at the junctures connecting the bottom and sides were investigated for corner radii from 0 to 8.5 percent of the body width. At low subsonic Mach numbers (free stream Mach number approximately equal 0.3) the aerodynamic characteristics are affected significantly by changes in Reynolds number and body corner radius. With increase in Mach number to free stream Mach number = 0.9 the effect of Reynolds number seems to vanish, but a significant effect of body corner radius remains.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6615 , A-4153
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  • 167
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    Tolerance of Mach 2.50 axisymmetric mixed-compression inlets to upstream flow variations (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation of the tolerances of two Mach 2.50 axisymmetric mixed-compression inlets to upstream flow variations was conducted. Tolerances of each inlet to angle of attack as a function of decreasing free-stream Mach number were obtained. A local region of overcompression was formed on the leeward side of the inlet at maximum angle of attack before unstart. This region of overcompression corresponded to local subsonic flow conditions ahead of the geometric throat. A uniform Mach number gradient of 0.10 at the cowl lip plane did not affect the inlet's pressure recovery, mass flow ratio, or diffuser exit total-pressure distortion.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2433 , E-6452
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  • 168
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    Splitting supersonic nozzle flow into separate jets by overexpansion into a multilobed divergent nozzle (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Air flowing from a convergent nozzle at pressure ratios greater than 2.5 has been split into eight separate jets by overexpansion of the flow into a divergent, eight-lobed passage. The splitting of the flow is accompanied by a decrease in the nozzle axial centerline Mach number. This in part is due to the radial inflow of secondary air between the lobes toward the nozzle centerline. Each of the smaller jets is partially split after it leaves the end of the divergent lobed section of the nozzle, thus creating a velocity profile having 16 peaks. At a pressure ratio of 3.5 the flow decelerates to Mach 1 in three convergent nozzle throat diameters. Convergent nozzle flow normally requires 12 diameters to reach Mach 1. The nozzle has a sound attenuation of 12 decibels with a thrust loss of 9 percent for the best configuration tested.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6667 , E-6640
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  • 169
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    FORTRAN program for calculating velocities in the meridional plane of a turbomachine 1: Centrifugal compressor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The program will determine the velocities in the meridional plane of a backward-swept impeller, a radial impeller, and a vaned diffuser. The velocity gradient equation with the assumption of a hub-to-shroud mean stream surface is solved along arbitrary quasi-orthogonals in the meridional plane. These quasi-orthogonals are fixed straight lines.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6701 , E-6592
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  • 170
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    Inverse problem approach to the design of short two-dimensional diffusers. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Keywords: AERODYNAMICS
    Type: ASME PAPER 72-WA/GT-6 , Winter Annual Meeting of the American Society of Mechanical Engineers; Nov 26, 1972 - Nov 30, 1972; New York, NY
    Format: text
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  • 171
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    Nacelle-airframe interference at low supersonic Mach numbers. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The aerodynamic interference between the propulsion system and airframe for a low supersonic transport with wing-mounted nacelles is examined. Both a flowfield analysis and the equivalent body approach were used to predict the interference lift, drag, and pitching moment as functions of nacelle size, shape, and position. The results indicate that the interference lift and pitching moment, as well as drag, must be included in the analysis to properly assess the interference effects. In addition, the performance of the basic wing was found to play an important role in determining the effectiveness of the interference lift in reducing the net installation drag. Based on a conservative prediction, the interference effects can reduce the installed propulsion system drag to 40% of the isolated drag of the nacelles. Furthermore, including the interference effects in the optimization of the engine cycle from a thermodynamic and weight standpoint can result in a considerable reduction in the net propulsion system weight fraction (fuel plus engines) while increasing the optimum engine bypass ratio of a typical transport vehicle.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-1113 , Joint Propulsion Specialist Conference; Nov 29, 1972 - Dec 01, 1972; New Orleans, LA; US
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  • 172
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    New technology in turbine aerodynamics. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Cursory review of some recent work that has been done in turbine aerodynamic research. Topics discussed include the aerodynamic effect of turbine coolant, high work-factor (ratio of stage work to square of blade speed) turbines, and computer methods for turbine design and performance prediction. Experimental cooled-turbine aerodynamics programs using two-dimensional cascades, full annular cascades, and cold rotating turbine stage tests are discussed with some typical results presented. Analytically predicted results for cooled blade performance are compared to experimental results. The problems and some of the current programs associated with the use of very high work factors for fan-drive turbines of high-bypass-ratio engines are discussed. Computer programs have been developed for turbine design-point performance, off-design performance, supersonic blade profile design, and the calculation of channel velocities for subsonic and transonic flowfields. The use of these programs for the design and analysis of axial and radial turbines is discussed.
    Keywords: AERODYNAMICS
    Type: Turbomachinery Symposium; Oct 24, 1972 - Oct 26, 1972; College Station, TX
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  • 173
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    Flight and wind tunnel investigation of the effects of Reynolds number on installed boattail drag at subsonic speeds (1972)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A flight and wind tunnel investigation was conducted to determine the effects of Reynolds number on the installed boattail drag of an underwing nacelle. Tests were run on a modified F-106B aircraft and 0.05 and 0.22 scale wind tunnel models. Tests were conducted at Mach numbers of 0.6 and 0.9 and over a 16 to 1 range of Reynolds numbers. Highest drag was obtained at intermediate Reynolds numbers corresponding to about the lowest flight values and that of the 0.22 scale model. Significantly lower drag was obtained at both higher and lower Reynolds numbers.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-68162 , E-7221 , Aerospace Sci. Meeting and Tech. Display; Jan 10, 1973 - Jan 12, 1973; Washington, D. C.; United States
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  • 174
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    Investigation of the stability of the tip vortex generated by hovering propellers and rotors. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The objective of this experimental and theoretical investigation was to determine what factors and mechanisms are involved in vortex interaction and instability and how these phenomena manifest themselves. To answer these questions, the schlieren method of flow visualization was used to observe the wakes generated by two- and four-bladed model propellers and rotors. A concurrent free-wake analysis was conducted for comparative purposes. Schlieren pictures showing wake asymmetry, interaction, and instability are presented. Various factors and mechanisms believed to be responsible for these are discussed along with the effects produced by the number of blades, collective pitch, and tip speed. Free-wake calculations that qualitatively confirm those factors responsible for wake asymmetry and interaction are also presented.
    Keywords: AERODYNAMICS
    Type: Atmospheric Flight Mechanics Conference; Sep 11, 1972 - Sep 13, 1972; Palo Alto, CA*Moffett Field, CA
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  • 175
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    Forward flight effects on mixer nozzle design and noise considerations for STOL externally blown flap systems. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Experimental data of the peak axial-velocity decay in a moving airstream are presented for several types of nozzles. The nozzles include a six-tube mixer nozzle of a type considered for reduction of jet-flap interaction noise for externally-blown-flap STOL aircraft. The effect of secondary flow on the core flow velocity decay of a bypass nozzle is also discussed. Tentative correlation equations are suggested for the configurations evaluated. Recommendations for minimizing forward velocity effects on velocity decay and jet-flap interaction noise are made.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-792 , American Institute of Aeronautics and Astronautics, Aircraft Design, Flight Test, and Operations Meeting; Aug 07, 1972 - Aug 09, 1972; Los Angeles, CA
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  • 176
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    Compressible boundary layer with normal pressure gradients: Quasi-similarity equations - Their properties at the wall and at sharp and blunt leading edges. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The failure of most viscous-inviscid interaction methods at strong interactions is attributed to the presence of a normal pressure gradient. A new theory is proposed for supersonic laminar boundary layers that can generate normal pressure gradients. The Navier-Stokes equations are reexamined by an order of magnitude analysis and all first and second order terms are retained. The approximation is found to be dependent not only on the boundary layer thickness but also on the ratio of the dimensionless viscosity and density. The equations are transformed into two quasi-similar, nonlinear, third order, ordinary integro-differential equations for the velocity and pressure as functions of a single transverse variable. The properties of the equations at the boundaries are discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-696 , American Institute of Aeronautics and Astronautics, Fluid and Plasma Dynamics Conference; Jun 26, 1972 - Jun 28, 1972; Boston, MA
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  • 177
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    Analytic prediction of dynamic stall characteristics. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A new approach is presented for solving the problem of predicting dynamic stall characteristics. In this approach, the unsteady aerodynamic characteristics are related theoretically to static aerodynamic characteristics which are readily available for a great number of airfoil shapes. Using these static experimental data as an input, the developed analytical method predicts dynamic stall characteristics that are in good agreement with available experimental data. The analysis is also extended to include frequency and amplitude modulation effects.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-682 , American Institute of Aeronautics and Astronautics, Fluid and Plasma Dynamics Conference; Jun 26, 1972 - Jun 28, 1972; Boston, MA
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  • 178
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    Three dimensional nozzle-exhaust flow field analysis by a reference plane technique. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A numerical method based on reference plane characteristics has been developed for the calculation of highly complex supersonic nozzle-exhaust flow fields. The difference equations have been developed for three coordinate systems. Local reference plane orientations are employed using the three coordinate systems concurrently thus catering to a wide class of flow geometries. Discontinuities such as the underexpansion shock and contact surfaces are computed explicitly for nonuniform vehicle external flows. The nozzles considered may have irregular cross-sections with swept throats and may be stacked in modules using the vehicle undersurface for additional expansion. Results are presented for several nozzle configurations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-704 , American Institute of Aeronautics and Astronautics, Fluid and Plasma Dynamics Conference; Jun 26, 1972 - Jun 28, 1972; Boston, MA
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  • 179
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    A finite-difference method for transonic airfoil design. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: This paper describes an inverse method for designing transonic airfoil sections or for modifying existing profiles. Mixed finite-difference procedures are applied to the equations of transonic small disturbance theory to determine the airfoil shape corresponding to a given surface pressure distribution. The equations are solved for the velocity components in the physical domain and flows with embedded shock waves can be calculated. To facilitate airfoil design, the method allows alternating between inverse and direct calculations to obtain a profile shape that satisfies given geometric constraints. Examples are shown of the application of the technique to improve the performance of several lifting airfoil sections. The extension of the method to three dimensions for designing supercritical wings is also indicated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-679 , American Institute of Aeronautics and Astronautics, Fluid and Plasma Dynamics Conference; Jun 26, 1972 - Jun 28, 1972; Boston, MA
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  • 180
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    Numerical computation of multishocked, three-dimensional supersonic flow fields with real gas effects. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A computational procedure is presented which is capable of determining the supersonic flow field surrounding three-dimensional wing-body configurations such as a delta-wing space shuttle. The governing equations in conservation-law form are solved by a finite difference method using a second-order noncentered algorithm between the body and the outermost shock wave, which is treated as a sharp discontinuity. Secondary shocks which form between these boundaries are captured automatically, and the intersection of these shocks with the bow shock posed no difficulty. Resulting flow fields about typical blunt nose shuttle-like configurations at angle of attack are presented. The differences between perfect and real gas effects for high Mach number flows are shown.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-702 , American Institute of Aeronautics and Astronautics, Fluid and Plasma Dynamics Conference; Jun 26, 1972 - Jun 28, 1972; Boston, MA
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  • 181
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    An experimental investigation of base heating on typical Mars entry body shapes. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Hypersonic experiments on wire supported high angle blunt cones were conducted to determine the base heating characteristics. Heating distributions were obtained for several base configurations with variations in free stream Reynolds Number and gas composition. The dependence of the heating in the base region on angle of attack was also investigated. It was found that gas composition effects can be accounted for by comparison at equal Reynolds Number. Angle of attack effects can result in either increasing or decreasing base heating depending upon the location in the base region.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 72-317 , American Institute of Aeronautics and Astronautics, Thermophysics Conference; Apr 10, 1972 - Apr 12, 1972; San Antonio, TX
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  • 182
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    TRW vortex-lattice method subsonic aerodynamic analysis for multiple-lifting-surfaces (N. surface) TRW program number HA010B (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The program was designed to provide solutions of engineering accuracy for determining the aerodynamic loads on single- or multiple-lifting-surface configurations that represent vehicles in subsonic flight, e.g., wings, wing-tail, wing-canard, lifting bodies, etc. The preparation is described of the input data, associated input arrangement, and the output format for the program data, including specification of the various operational details of the program such as array sizes, tape numbers utilized, and program dumps. A full description of the underlying theory used in the program development and a review of the program qualification tests are included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-128588 , TRW-20029-H110-RO-00
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  • 183
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    Additional three-dimensional boundary-layer computations for a finite swept wing (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Calculations have been made of the three-dimensional, compressible, turbulent boundary layer on the finite supercritical wing of the NASA modified F-8 transonic research airplane. The calculations were based on the wing pressure distribution measured in flight at M = 0.90, instead of on wind tunnel data at M = 0.50 and 0.99. Data on the boundary-layer thickness, displacement thickness, skin-friction components, and integrated streamwise skin friction are presented for points along the streamwise stations at which pressure measurements were made.
    Keywords: AERODYNAMICS
    Type: NASA-CR-132335
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  • 184
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    A method for calculating aerodynamic heating on sounding rocket tangent ogive noses (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A method is presented for calculating the aerodynamic heating and shear stresses at the wall for tangent ogive noses that are slender enough to maintain an attached nose shock through that portion of flight during which heat transfer from the boundary layer to the wall is significant. The lower entropy of the attached nose shock combined with the inclusion of the streamwise pressure gradient yields a reasonable estimate of the actual flow conditions. Both laminar and turbulent boundary layers are examined and an approximation of the effects of (up to) moderate angles-of-attack is included in the analysis. The analytical method has been programmed in FORTRAN 4 for an IBM 360/91 computer.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-69896
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  • 185
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    Effects of vane-induced rotation on diffuser flow distortion in an axisymmetric mixed-compression inlet (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation of vane-induced flow rotation to modify distorted steady-state total-pressure patterns in the subsonic diffuser of a supersonic mixed-compression inlet was conducted. Radial static-pressure gradients generated by the rotation was the mechanism used to modify the total-pressure distributions. Significant redistribution of circumferential distortion patterns into more compatible radial patterns was realized, but flow problems near the duct walls reduced the general effectiveness of the technique. Total-pressure losses associated with the swirl vanes were slight. Limited turbulence data indicated that vane istallation resulted in reduced turbulence levels.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2752 , E-7245
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  • 186
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    A pertinent solution of helicopter rotor flapping stability (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The stability of the flapping motion of a single blade of a helicopter rotor is examined using the techniques of perturbation theory. The equation of motion studied is linear, with periodic aerodynamic coefficients due to the forward speed of the rotor. Solutions are found for four cases: small and large advance ratio and small and large Lock number. The perturbation techniques appropriate to each case are discussed and illustrated in the course of the analysis. The application of perturbation techniques to other problems in rotor dynamics is discussed. It is concluded that perturbation theory is a powerful mathematical technique which should prove very useful in analyzing some of the problems of helicopter dynamics.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62165
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  • 187
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    Numerical method for design of minimum-drag supersonic wing camber with constraints on pitching moment and surface deformation (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A numerical method, based on linearized theory, for designing minimum-drag supersonic wing camber surfaces of arbitrary planform for a given lift, with options for constraining the pitching moment and/or the surface deformation at the trailing edge of the root chord and for selecting any desired combination of eight specified wing-loading distributions to be employed in the optimization procedure is presented. Two examples are given to illustrate applications of the method. The results indicate that relatively small drag penalties are incurred in designing wings to be self-trimming and to have a reasonable camber surface.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-7097 , L-8585
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  • 188
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    Performance of a single-stage axial-flow transonic compressor stage with a blade tip solidity of 1.7 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The overall and blade-element performance of a transonic compressor stage is presented over the stable operating range at rotative speeds from 50 to 100 percent of design speed. Stage peak efficiency of 0.784 was obtained at a weight flow of 28.6 kilograms per second and a pressure ratio of 1.706. Stall margin at design speed was 11.4 percent. The peak efficiency being significantly less than design efficiency was attributed to: (1) the stator loss and the radial gradient of losses being much higher than design, (2) the losses and blockages associated with the rotor part-span dampers not being incorporated into the design, and (3) mismatch of the rotor and stator badle elements.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2658 , E-6730
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  • 189
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    Leading-edge force features of the aerodynamic finite element method. (1972)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: Description of a practical procedure for computing the wing leading-edge thrust distribution by the finite element method. When incorporated into a wing-body aerodynamic computer program, the technique is capable of predicting (at subsonic and supersonic speeds) the leading-edge thrust distribution (and therefore, the lateral-directional stability derivatives due to roll) and the nonlinear aerodynamic characteristics of low aspect-ratio wings with leading edge separation due to the application of suction technology.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 9; Dec. 197
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  • 190
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    Slowly oscillating lifting surfaces at subsonic and supersonic speeds. (1972)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: The paper presents an unsteady aerodynamic influence coefficient method based on the low-frequency approximation. The influence coefficients are of a type which have been used to compute steady flow about wing-body combinations; therefore, the new method may be extended readily to low-frequency unsteady flow about wing-body combinations. The validity of the method is demonstrated by comparisons with numerical results from conventional, unsteady lifting surface methods. The method is valid for arbitrary wings in supersonic flow and for wings of finite span in subsonic flow. The method, when extended to include wing-body-tail interactions, will have important applications for predicting stability, control, and gust response characteristics of large airplanes. Dynamic stability derivatives and pressure distributions are given for several planforms. The comparison with either analytical or other well established numerical methods shows good agreement.-
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 9; Nov. 197
    Format: text
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  • 191
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    On the mechanism of dynamic stall. (1972)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: The sequence of events comprising dynamic stall of an airfoil is discussed, with emphasis on the role of the leading edge laminar separation bubble and shed vortex. A simple bubble model, based on a combination of theoretical and experimental investigations, is used to discuss the events prior to the shedding of the vortex, and provides the basis for a heuristic estimate of the delay in the occurrence of dynamic stall on a pitching airfoil. The evidence for the existence and dominant effect of the leading edge vortex on the dynamic stall required (but in most cases not presently available) for the prediction of the effects of stall on helicopter rotor blades are discussed. It is the intention of this paper to focus attention on the laminar separation bubble and the shed leading edge vortex as the dominant features of the dynamic stall mechanism in the hope of stimulating greater emphasis on these features in future dynamic stall research.
    Keywords: AERODYNAMICS
    Type: AD-762973 , AROD-4846-15-E , American Helicopter Society; vol. 17
    Format: text
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  • 192
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    Hypersonic, viscous shock layer with chemical nonequilibrium for spherically blunted cones. (1972)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 10; Oct. 197
    Format: text
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  • 193
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    An experimental investigation of an oblique-wing and body combination at Mach numbers between 0.60 and 1.40 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation was conducted in an 11- by 11-foot wind tunnel to determine the aerodynamic characteristics of an oblique high aspect ratio wing in combination with a high fineness-ratio Sears-Haack body. Longitudinal and lateral-directional stability data were obtained at wing yaw angles from 0 deg to 60 deg over a test Mach number range from 0.6 to 1.4 for angles of attack between minus 6 deg and 9 deg. The effects of changes in Reynolds number, dihedral, and trailing-edge angle were studied along with the effects of a roughness strip on the upper and lower surfaces of the wing. Flow-visualization studies were made to determine the nature of the flow on the wing surfaces.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62207
    Format: application/pdf
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  • 194
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    Study of induced load and stress, volume 3 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An analytical and experimental investigation into the effects of blade tip clearance on inducer performance and of leading edge sweepback on both blade pressure loading and performance was performed. Tip clearance flow was represented with a vortex flow model and measured data from previous inducer tests at three clearances were correlated with model predictions. A leading edge model was added to an existing inducer internal flow analysis, tests with two sweepbacks were conducted, and blade pressure and performance predictions were correlated with measured data.
    Keywords: AERODYNAMICS
    Type: NASA-CR-72712 , PWA-FR-3704-VOL-3
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  • 195
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    Low-speed aerodynamic characteristics of an airfoil optimized for maximum lift coefficient (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation has been conducted in the Langley low-turbulence pressure tunnel to determine the two-dimensional characteristics of an airfoil optimized for maximum lift coefficient. The design maximum lift coefficient was 2.1 at a Reynolds number of 9.7 million. The airfoil with a smooth surface and with surface roughness was tested at angles of attack from 6 deg to 26 deg, Reynolds numbers (based on airfoil chord) from 2.0 million to 12.9 million, and Mach numbers from 0.10 to 0.35. The experimental results are compared with values predicted by theory. The experimental pressure distributions observed at angles of attack up to at least 12 deg were similar to the theoretical values except for a slight increase in the experimental upper-surface pressure coefficients forward of 26 percent chord and a more severe gradient just behind the minimum-pressure-coefficient location. The maximum lift coefficients were measured with the model surface smooth and, depending on test conditions, varied from 1.5 to 1.6 whereas the design value was 2.1.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-7071 , L-8491
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  • 196
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    Atmospheric Flight Mechanics Conference, 2nd, Palo Alto and Moffett Field, Calif., September 11-13, 1972, Informal Papers. (1972)
    In:  Other Sources
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    Publication Date: 2019-07-27
    Description: Limitations in the acquisition of nonlinear aerodynamic coefficients from free-oscillation data by means of the Chapman-Kirk technique, SAM-D control test vehicle trajectory plannning and flight test analysis, and determination of aerodynamic drag from radar data are among the topics covered in papers concerned with atmospheric flight mechanics. Other areas covered include fixed and rotary-wing aircraft, ordnance and reentry vehicles, and analysis and measurement techniques. Individual items are announced in this issue.
    Keywords: AERODYNAMICS
    Type: Atmospheric Flight Mechanics Conference; Sept. 11-13, 1972; Palo Alto, CA*Moffett Field, CA
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  • 197
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    Supercritical aerodynamics - Worthwhile over a range of speeds. (1972)
    In:  Other Sources
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    Publication Date: 2019-08-13
    Description: Supercritical flow exists whenever a high enough forward speed causes local flow over a lifting surface or body to exceed the sonic or critical value. The principal difference between conventional subsonic aerodynamic technology and supercritical technology lies in the cross-sectional profile of lifting surfaces. The characteristics of the supercritical airfoil suggests three potential benefits from applications to civil aircraft. For aircraft designed to operate at moderate subsonic speeds the supercritical airfoil may permit the reduction of structural weight. Supercritical technology would have a second application in permitting efficient high subsonic speed cruise by delaying the transonic drag rise. Another advantage of the thick supercritical wing shows up at low speeds.
    Keywords: AERODYNAMICS
    Type: Astronautics and Aeronautics; 10; Aug. 197
    Format: text
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  • 198
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    Aeronautical Engineering: A special bibliography with indexes, supplement 24, November 1972 (1972)
    In:  CASI
    Details
    Publication Date: 2019-08-27
    Description: This special bibliography lists 399 reports, articles, and other documents introduced into the NASA scientific and technical information system in October 1972.
    Keywords: AERODYNAMICS
    Type: NASA-SP-7037(24)
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  • 199
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    Aeronautical engineering: A special bibliography with indexes, supplement 20 (1972)
    In:  CASI
    Details
    Publication Date: 2019-08-27
    Description: This special bibliography lists 437 reports, articles, and other documents introduced into the NASA scientific and technical information system in June 1972.
    Keywords: AERODYNAMICS
    Type: NASA-SP-7037(20)
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  • 200
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    Aeronautical Engineering: A special bibliography, supplement 25 December 1972 (1972)
    In:  CASI
    Details
    Publication Date: 2019-08-27
    Description: This special bibliography lists 373 reports, articles, and other documents introduced into the NASA scientific and technical information system in November 1972.
    Keywords: AERODYNAMICS
    Type: NASA-SP-7037(25)
    Format: application/pdf
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