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  • General Chemistry  (524)
  • Cell & Developmental Biology  (107)
  • Aerodynamics  (53)
  • Fluid Mechanics and Heat Transfer  (25)
  • 1960-1964  (709)
  • 1960  (709)
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  • 1960-1964  (709)
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  • 1
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    A Method of Reducing Aeroelastic Effects of Highly Swept Wings (1960)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: The greatest efficiency for a lifting surface at supersonic speeds, according to the theoretical considerations of reference 1, can be attained if the leading edge is swept well behind the Mach cone and the highest aspect ratio which is structurally possible is employed. Such a wing, designed for a Mach number of 3.0, would have 80 deg. of sweepback. Aeroelastic effects have 〈 been shown 3 to be considerable for a wing with 60deg of sweepback and designed for a Mach number of 2.0. The wing shown was found theoretically to have considerable loss in maximum lift-drag ratio attributable to aeroelasticity. This wing has 12-per cent-thick Clark-Y airfoils normal to the wing leading edge. If it were of solid aluminum and flying at a dynamic pressure of 2,400 lbs./sq.ft. (flexibility parameter qb(exp. 4) /El(0) = 7.8), analysis indicates that the wing would deflect so as to reduce the maximum lift-drag ratio about 30 per cent.
    Keywords: Aerodynamics
    Type: Journal of the Aerospace Sciences; Volume 27; No. 8; 634-635
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  • 2
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    Measurements of Skin Friction of the Compressible Turbulent Boundary Layer on a Cone with Foreign Gas Injection (1960)
    In:  Other Sources
    Details
    Publication Date: 2018-06-05
    Description: Measurements of average skin friction of the turbulent boundary layer have been made on a 15deg total included angle cone with foreign gas injection. Measurements of total skin-friction drag were obtained at free-stream Mach numbers of 0.3, 0.7, 3.5, and 4.7 and within a Reynolds number range from 0.9 x 10(exp 6) to 5.9 x 10(exp 6) with injection of helium, air, and Freon-12 (CCl2F2) through the porous wall. Substantial reductions in skin friction are realized with gas injection within the range of Mach numbers of this test. The relative reduction in skin friction is in accordance with theory-that is, the light gases are most effective when compared on a mass flow basis. There is a marked effect of Mach number on the reduction of average skin friction; this effect is not shown by the available theories. Limited transition location measurements indicate that the boundary layer does not fully trip with gas injection but that the transition point approaches a forward limit with increasing injection. The variation of the skin-friction coefficient, for the lower injection rates with natural transition, is dependent on the flow Reynolds number and type of injected gas; and at the high injection rates the skin friction is in fair agreement with the turbulent boundary layer results.
    Keywords: Aerodynamics
    Type: Journal of Aerospace Sciences; Volume 27; No. 5; 321-333
    Format: text
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  • 3
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    Effects of Fixing Boundary-Layer Transition for a Swept- and a Triangular-Wing and Body Combination at Mach Numbers from 0.60 to 1.40 (1960)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The aerodynamic effects of fixing boundary-layer transition for a swept- and a triangular-wing configuration have been determined from tests of two small-scale wing-body models. The wings had an aspect ratio of 2.99 and 3-percent-thick biconvex sections. Lift, pitching-moment, and drag data were obtained at Mach numbers ranging from 0.60 to 1.40 for angles of attack between -2 deg and about 15 deg. The Reynolds number of the tests was generally 1.5 million; however, minimum drag measurements were made for both models over a range of Reynolds numbers from 1.0 million to about 3.0 or 4.0 million.
    Keywords: Aerodynamics
    Type: NASA-TN-D-312
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  • 4
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    Ground Effect for Lifting Rotors in Forward Flight (1960)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A theoretical analysis indicates that, for rotors, ground effect decreases rapidly with increases in either height above the ground or forward speed. The decrease with height above the ground in forward flights is greater than that in hovering. The major part of the decrease in ground effect with forward speed occurs at speeds less than 1.5 times the hovering mean induced velocity. Consequently, the total induced velocity at the rotor center increases rather than decreases when a helicopter gathers speed at low height above the ground.
    Keywords: Aerodynamics
    Type: NASA-TN-D-234
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  • 5
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    Effect of Ground Proximity on Aerodynamic Characteristics of Two Horizontal-Attitude Jet Vertical-Take-Off-and-Landing Airplane Models (1960)
    In:  CASI
    Details
    Publication Date: 2019-06-25
    Description: An investigation has been made to study the effect of ground proximity on the aerodynamic characteristics of two jet vertical-take-off-and-landing airplane models in which the fuselage remains in a horizontal attitude for the take-off and landing. The first model (called the tilt-wing model) had a tilting wing-engine assembly which was set at 90 deg incidence for the take-off and landing. The second model, called the deflected-jet model) had a cascade of retractable turning vanes to deflect the exhaust of the horizontally mounted jet engines downward for vertical take-off and landing while the entire model remained in a horizontal attitude. With the models at various heights above the ground in the take-off and landing configuration, the lift, drag, and pitching moment were measured and tuft surveys were made to determine the flow field caused by the jet exhaust. The tilt-wing model experienced a loss of lift of less than 3 percent near the ground. The deflected-jet model, however, suffered losses in lift as high as 45 percent near the ground because of a low pressure region under the model caused by the entrainment of air by the jet exhaust as it spread out along the ground. This loss in lift for the deflected-jet configuration could probably be reduced to less than 5 percent by the use of a longer landing gear and a high wing location.
    Keywords: Aerodynamics
    Type: NASA-TN-D-419 , L-1059
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  • 6
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    Tests of Vortex-Ring Parachute at Supersonic Speed in the Langley Unitary Plan Wind Tunnel (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: For the test, the 12-inch-diameter "Vortex-Ring" parachute was towed behind a conical-nosed cylindrical body 2.25 inches in diameter. The tow-cable length was 24 inches, and was attached to the cylindrical body through a large swivel and to the parachute through a smaller swivel. The attachment between the large swivel an the cylindrical body failed after about 1 minute's operation. Mach number was approximately 2.2, dynamic pressure was approximately 150 pounds per square foot, and camera speed was approximately 3000 frames per second.
    Keywords: Aerodynamics
    Type: L-560
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  • 7
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    Aerodynamic Characteristics at Low Speed of a Reentry Configuration Having Rigid Retractable Conical Lifting Surfaces (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An exploratory investigation has been made in the Langley 300 MPH 7 by 10 foot tunnel to study the low-speed static longitudinal and lateral stability characteristics of a reentry configuration having rigid retractable conical lifting surfaces that unfolded from the surface of a conical fuselage. The model also had curved tail surfaces that unfolded from a cylindrical aft section attached to the cone. Longitudinal tests were made through an angle-of-attack range from -4 deg to 90 deg and limited lateral tests were made through an angle-of-sideslip range from -12 deg to 32 deg at an angle of attack of 0 deg. The tail surface provided longitudinal trim to maximum lift and beyond and up to an angle of attack of 51 deg for a center-of-moment location of 42.9 percent mean aerodynamic chord. For this center-of-moment position the model had a static margin of 12 percent mean aerodynamic chord at the lower lift coefficients and was longitudinally stable up to a lift coefficient between 1.0 and 1.2. Neutral stability occurred from lift coefficient of 1.0 up to near maximum lift coefficient. The maximum value of trimmed lift-drag ratio was 4.85 at a lift coefficient of approximately 0.3 and a trimmed angle of attack of approximately 10 deg. The configuration was directionally stable throughout the test angle of sideslip range for an angle of attack of 0 deg.
    Keywords: Aerodynamics
    Type: NASA-TN-D-622 , L-1180
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  • 8
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    A Technique for Determining Relaxation Times by Free-Flight Tests of Low-Fineness-Ratio Cones; with Experimental Results for Air at Equilibrium Temperatures up to 3440 K (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: This report describes a technique which combines theory and experiments for determining relaxation times in gases. The technique is based on the measurement of shapes of the bow shock waves of low-fineness-ratio cones fired from high-velocity guns. The theory presented in the report provides a means by which shadowgraph data showing the bow waves can be analyzed so as to furnish effective relaxation times. Relaxation times in air were obtained by this technique and the results have been compared with values estimated from shock tube measurements in pure oxygen and nitrogen. The tests were made at velocities ranging from 4600 to 12,000 feet per second corresponding to equilibrium temperatures from 35900 R (19900 K) to 6200 R (34400 K), under which conditions, at all but the highest temperatures, the effective relaxation times were determined primarily by the relaxation time for oxygen and nitrogen vibrations.
    Keywords: Aerodynamics
    Type: NASA-TN-D-327
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  • 9
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    Photographic Evidence of Streamwise Arrays of Vortices in Boundary-Layer Flow (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Photographs are presented of various models coated with fluorescent oil to show evidence of surface vortices at a Mach number of 3.03. Vortex formation was evidently present on models with forward-facing steps, rearward-facing steps, wires, discrete surface particles, or unswept flat surfaces with sharp leading edges. Some photographs are also presented for the models coated with a sublimation material which clearly indicates the location of boundary-layer transition; however, it does not show the vortices as clearly as the fluorescent oil. The study was made on the models at an angle of attack of 0 deg at unit Reynolds numbers of 7.7 and 10.7 million per foot. The spacing of the vortices as indicated by the flow studies on the unswept model was smaller at the higher Reynolds number in accordance with Gortler's theory. The flow studies also indicated that stable surface vortices produced by either steps or surface roughness persisted over model areas known to have turbulent boundary layers.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-328
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  • 10
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    Wind-Tunnel Investigation of a Small-Scale Model of an Aerial Vehicle Supported by Tilting Ducted Fans (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A wind-tunnel investigation has been made to study the static longitudinal and lateral stability characteristics of a simplified aerial vehicle supported by ducted fans that tilt relative to the airframe. The ducts were in a triangular arrangement with one duct in front and two at the rear in order to minimize the influence of the downwash of the front duct on the rear ducts. The results of the investigation were compared with those of a similar investigation for a tandem two-duct arrangement in which the ducts were fixed (rather than tiltable) relative to the airframe, since the three-duct configuration had been devised in an attempt to avoid some of the deficiencies of the tandem fixed-duct configuration. The results of the investigation indicated that the tilting-duct arrangement had less noseup pitching moment for a given forward speed than the tandem fixed-duct arrangement. The model had less angle-of-attack instability than the tandem fixed-duct arrangement. The model was directionally unstable but had a positive dihedral effect throughout the test speed range.
    Keywords: Aerodynamics
    Type: NASA-TN-D-409 , L-961
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  • 11
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    The Longitudinal Aerodynamic Characteristics of a Sweptback Wing-Body Combination With and Without End Plates at Mach Numbers from 0.40 to 0.93 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation was made at high subsonic speeds in the Langley high-speed 7- by 10-foot tunnel to determine the effect of end plates on the longitudinal aerodynamic characteristics of a sweptback wing-body combination with and without drooped chord-extensions. The wing had 45 deg sweepback of the quarter-chord line, an aspect ratio of 4, a taper ratio of 0.3, and NACA 65AO06 airfoil sections parallel to the plane of symmetry, and was mounted near the rear of a body of revolution having a fineness ratio of approximately 8. The results indicated that the addition of the end plates to either the wing with drooped chord-extensions or to the wing without drooped chord-extensions slightly increased the lift in the low angle-of-attack range but slightly decreased the lift at moderate and high angles of attack. The addition of the end plates to the wing without the chord-extensions caused a small increase in the maximum lift-drag ratio at Mach numbers below 0.65 and a slight decrease at the higher Mach numbers; however, for the addition of the end plates to the wing with the chord- extensions the maximum lift-drag ratio was slightly decreased below a Mach number of 0.88, while a slight increase occurred for the higher Mach numbers. The addition of the end plates to the wings with and without the chord-extensions caused the static longitudinal stability to increase considerably for all Mach numbers; however, only a slight reduction in the aerodynamic-center variation with Mach number was observed.
    Keywords: Aerodynamics
    Type: NASA-TN-D-389 , L-834
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  • 12
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    Some Examples of the Applications of the Transonic and Supersonic Area Rules to the Prediction of Wave Drag (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The experimental wave drags of bodies and wing-body combinations over a wide range of Mach numbers are compared with the computed drags utilizing a 24-term Fourier series application of the supersonic area rule and with the results of equivalent-body tests. The results indicate that the equivalent-body technique provides a good method for predicting the wave drag of certain wing-body combinations at and below a Mach number of 1. At Mach numbers greater than 1, the equivalent-body wave drags can be misleading. The wave drags computed using the supersonic area rule are shown to be in best agreement with the experimental results for configurations employing the thinnest wings. The wave drags for the bodies of revolution presented in this report are predicted to a greater degree of accuracy by using the frontal projections of oblique areas than by using normal areas. A rapid method of computing wing area distributions and area-distribution slopes is given in an appendix.
    Keywords: Aerodynamics
    Type: NASA-TN-D-446 , L-1000
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  • 13
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    Heat Transfer to 36.75 and 45 degree Swept Blunt Leading Edges in Free Flight at Mach Numbers from 1.70 to 2.99 and From 2.50 to 4.05 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A flight investigation has been conducted to study the heat transfer to swept-wing leading edges. A rocket-powered model was used for the investigation and provided data for Mach number ranges of 1.78 to 2.99 and 2.50 to 4.05 with corresponding free-stream Reynolds number per foot ranges of 13.32 x 10(exp 6) to 19.90 x 10(exp 6) and 2.85 x 10(exp 6) to 4.55 x 10(exp 6). The leading edges employed were cylindrically blunted wedges ', three of which were swept 450 with leading-edge diameters of 1/4, 1/2, and 3/4 inch and one swept 36-750 with a leading-edge diameter of 1/2 inch. In the high Reynolds number range, measured values of heat transfer were found to be much higher than those predicted by laminar theory and at the larger values of leading-edge diameter were approaching the values predicted by turbulent theory. For the low Reynolds number range a comparison between measured and theoretical heat transfer showed that increasing the leading-edge diameter resulted in turbulent flow on the cylindrical portion of the leading edge.
    Keywords: Aerodynamics
    Type: NASA-TM-X-208
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  • 14
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    The Shock-Wave Patterns on a Cranked-Wing Configuration (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The shock-wave patterns of a complex configuration with cranked cruciform wings and a cone-cylinder body were examined to determine the interaction of the body bow wave with the flow field about the wing. Also of interest, was the interaction of the forward (760 sweptback) wing leading-edge wave with the rear (600 sweptback) wing leading-edge wave. The shadowgraph pictures of the model in free flight at a Mach number of 4.9, although not definitive, appear to indicate that the body bow wave crosses the outer wing panel after first being refracted either by the leading-edge wave of the 600 sweptback wing or by pressure fields in the flow crossing the wing.
    Keywords: Aerodynamics
    Type: NASA-TN-D-346 , A-433
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  • 15
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    Axial-Force Reduction by Interference Between Jet and Neighboring Afterbody (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Experimental results are presented for an exploratory investigation of the effectiveness of interference between jet and afterbody in reducing the axial force on an afterbody with a neighboring jet. In addition to the interference axial force., measurements are presented of the interference normal force and the center of pressure of the interference normal force. The free-stream Mach number was 2.94, the jet-exit Mach number was 2.71, and the Reynolds number was 0.25 x 10, based on body diameter. The variables investigated include static-pressure ratio of the jet (up to 9), nacelle position relative to afterbody, angle of attack (-5 deg to 10 deg), and afterbody shape. Two families of afterbody shapes were tested. One family consisted of tangent-ogive bodies of revolution with varying length and base areas. The other family was formed by taking a planar slice off a circular cylinder with varying angle between the plane and cylinder. The trends with these variables are shown for conditions near maximum jet-afterbody interference. The interference axial forces are large and favorable. For several configurations the total afterbody axial force is reduced to zero by the interference.
    Keywords: Aerodynamics
    Type: NASA-TN-D-332
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  • 16
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    Effects of Sweep Angle on the Boundary-Layer Stability Characteristics of an Untapered Wing at Low Speeds (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation was conducted in the Ames 12-Foot Low-Turbulence Pressure Tunnel to determine the effects of sweep on the boundary-layer stability characteristics of an untapered variable-sweep wing having an NACA 64(2)A015 section normal to the leading edge. Pressure distribution and transition were measured on the wing at low speeds at sweep angles of 0, 10, 20, 30, 40, and 50 deg. and at angles of attack from -3 to 3 deg. The investigation also included flow-visualization studies on the surface at sweep angles from 0 to 50 deg. and total pressure surveys in the boundary layer at a sweep angle of 30 deg. for angles of attack from -12 to 0 deg. It was found that sweep caused premature transition on the wing under certain conditions. This effect resulted from the formation of vortices in the boundary layer when a critical combination of sweep angle, pressure gradient, and stream Reynolds number was attained. A useful parameter in indicating the combined effect of these flow variables on vortex formation and on beginning transition is the crossflow Reynolds number. The critical values of crossflow Reynolds number for vortex formation found in this investigation range from about 135 to 190 and are in good agreement with those reported in previous investigations. The values of crossflow Reynolds number for beginning transitions were found to be between 190 and 260. For each condition (i.e., development of vortices and initiation of transition at a given location) the lower values in the specified ranges were obtained with a light coating of flow-visualization material on the surface. A method is presented for the rapid computation of crossflow Reynolds number on any swept surface for which the pressure distribution is known. From calculations based on this method, it was found that the maximum values of crossflow Reynolds number are attained under conditions of a strong pressure gradient and at a sweep angle of about 50 deg. Due to the primary dependence on pressure gradient, effects of sweep in causing premature transition are generally first encountered on the lower surfaces of wings operating at positive angles of attack.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-338
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  • 17
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    An Experimental Investigation and Correlation of the Heat Reduction to Nonporous Surfaces Behind a Porous Leading Edge Through Which Coolant is Ejected (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A configuration of a wing segment having constant chord thickness, 0 deg. sweep, a porous steel semicircular leading edge, and solid Inconel surfaces was tested in a Mach number 2.0 ethlyene-heated high-temperature air jet. Measurements were made of the wing surface temperatures at chordwise stations for several rates of helium flow through the porous leading edge. The investigation was conducted at stagnation temperatures ranging from 500 F to 2,400 F, at Reynolds numbers per foot ranging from 0.3 x 10(exp 7) to 1.2 x 10(exp 7), and at angles of attack of 0, +/- 5, and +/- 15 deg. The results indicated that the reduction of wing surface temperatures with respect to their values for no coolant flow, depended on the helium coolant flow rates and the distance behind the area of injection. The results were correlated in terms of the wall cooling parameter and the coolant flow-rate parameter, where the nondimensional flow rate was referenced to the cooled area up to the downstream position. For the same coolant flow rate, lower surface temperatures are achieved with a porous-wall cooling system. However, since flow-rate requirements decrease with increasing allowable surface temperatures, the higher allowable wall temperatures of the solid wall as compared to the structurally weaker porous wall- sharply reduce the flow-rate requirements of a downstream cooling system. Thus, for certain flight conditions it is possible to compensate for the lower efficiency of the downstream or solid-wall cooling system. For example, a downstream cooling system using solid walls that must be maintained at 1,800 F would require less coolant for Mach numbers up to 5.5 than would a porous-wall cooling system for which the walls must be maintained at temperatures less than or equal to 9000 F.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TM-X-235
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  • 18
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    Large-Scale Wind-Tunnel Tests and Evaluation of the Low-Speed Performance of a 35 deg Sweptback Wing Jet Transport Model Equipped with a Blowing Boundary-Layer-Control Flap and Leading-Edge Slat (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A wind-tunnel investigation was conducted to determine the effect of trailing-edge flaps with blowing-type boundary-layer control and leading-edge slats on the low-speed performance of a large-scale jet transport model with four engines and a 35 deg. sweptback wing of aspect ratio 7. Two spanwise extents and several deflections of the trailing-edge flap were tested. Results were obtained with a normal leading-edge and with full-span leading-edge slats. Three-component longitudinal force and moment data and boundary-layer-control flow requirements are presented. The test results are analyzed in terms of possible improvements in low-speed performance. The effect on performance of the source of boundary-layer-control air flow is considered in the analysis.
    Keywords: Aerodynamics
    Type: NASA-TN-D-333 , A-340
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  • 19
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    Evaluation of Blended Wing-Body Combinations with Curved Plan Forms at Mach Numbers Up to 3.50 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: This investigation is a continuation of the experimental and theoretical evaluation of the effects of wing plan-form variations on the aerodynamic performance characteristics of blended wing-body combinations. The present report compares previously tested straight-edged delta and arrow models which have leading-edge sweeps of 59.04 and 70-82 deg., respectively, with related models which have plan forms with curved leading and trailing edges designed to result in the same average sweeps in each case. All the models were symmetrical, without camber, and were generally similar having the same span, length, and aspect ratios. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local wing chords in a streamwise direction. The wing sections were computed by varying their shapes along with the body radii (blending process) to match the selected area distribution and the given plan form. The models were tested with transition fixed at Reynolds numbers of roughly 4,000,000 to 9,000,000, based on the mean aerodynamic chord of the wing. The characteristic effect of the wing curvature of the delta and arrow models was an increase at subsonic and transonic speeds in the lift-curve slopes which was partially reflected in increased maximum lift-drag ratios. Curved edges were not evaluated on a diamond plan form because a preliminary investigation indicated that the curvature considered would increase the supersonic zero-lift wave drag. However, after the test program was completed, a suitable modification for the diamond plan form was discovered. The analysis presented in the appendix indicates that large reductions in the zero-lift wave drag would be obtained at supersonic Mach numbers if the leading- and trailing-edge sweeps are made to differ by indenting the trailing edge and extending the root of the leading edge.
    Keywords: Aerodynamics
    Type: NASA-TM-X-379
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  • 20
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    A Wind-Tunnel Investigation of the Development of Lift on Wings in Accelerated Longitudinal Motion (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the development of lift on a wing during a simulated constant-acceleration catapult take-off. The investigation included models of a two-dimensional wing, an unswept wing having an aspect ratio of 6, a 35 deg. swept wing having an aspect ratio of 3.05, and a 60 deg. delta wing having an aspect ratio of 2.31. All the wings investigated developed at least 90 percent of their steady-state lift in the first 7 chord lengths of travel. The development of lift was essentially independent of the acceleration when based on chord lengths traveled, and was in qualitative agreement with theory.
    Keywords: Aerodynamics
    Type: NASA-TN-D-422 , L-1027
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  • 21
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    Transition Reynolds Numbers of Separated Flows at Supersonic Speeds (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Experimental research has been conducted on the effects of wall cooling, Mach number, and unit Reynolds number on the transition Reynolds number of cylindrical separated boundary layers on an ogive-cylinder model. Results were obtained from pressure and temperature measurements and shadowgraph observations. The maximum scope of measurements encompassed Mach numbers between 2.06 and 4.24, Reynolds numbers (based on length of separation) between 60,000 and 400,000, and ratios of wall temperature to adiabatic wall temperature between 0.35 and 1.0. Within the range of tile present tests, the transition Reynolds number was observed to decrease with increasing wall cooling, increase with increasing Mach number, and increase with increasing unit Reynolds number. The wall cooling effect was found to be four times as great when the attached boundary layer upstream of separation was cooled in conjunction with cooling of the separated boundary layer as when only the separated boundary layer was cooled. Wall cooling of both the attached and separated flow regions also caused, in some cases, reattachment in the otherwise separated region. Cavity resonance present in the separated region for some model configurations was accompanied by a large decrease in transition Reynolds number at the lower test Mach numbers.
    Keywords: Aerodynamics
    Type: NASA-TN-D-349 , A-178
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  • 22
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    Large-Scale Wind-Tunnel Tests of a Wingless Vertical Take-Off and Landing Aircraft: Preliminary Results (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Large-scale wind-tunnel tests were made of a wingless vertical take-off and landing aircraft at zero sideslip to determine performance and longitudinal stability and control characteristics at airspeeds from 0 to 70 knots. Roll control and rudder effectiveness were also obtained. Limitations in the propulsion system restricted the lift for which level flight could be simulated to approximately 1500 pounds. Test variables with roll control and rudder undeflected were airspeed, vane setting, angle of attack, elevator deflection, and power. In most of the tests angle of attack, elevator, and power were varied individually while the other four parameters were held constant at previously determined values required for simulating trimmed level flight. The majority of the tests were made with power on and tail on at airspeeds between 20 and 70 knots. However, a limited number of data were obtained for the following conditions: (1) at zero velocity, horizontal tail on, power on; (2) at forward velocity, tail off and power on; and (3) at forward velocity, tail on, but with power off.
    Keywords: Aerodynamics
    Type: NASA-TN-D-326
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  • 23
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    Transonic Aerodynamic Characteristics of a Model of a Proposed Six-Engine Hull-Type Seaplane Designed for Supersonic Flight (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.
    Keywords: Aerodynamics
    Type: NASA-TM-X-246
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  • 24
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    Effect on Gaseous Film Cooling of Coolant Injection Through Angled Slots and Normal Holes (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A study was made to determine the effect of coolant injection angularity on gaseous film-cooling effectiveness. In the correlation of experimental data an effective injection angle was defined by a vector summation of the coolant and mainstream gas flows. The cosine of this angle was used as a parameter to empirically develop a corrective term to qualify a correlating equation presented in Technical Note D-130 that was limited to tangential injection of the coolant. Data were also obtained for coolant injection through rows of holes normal to the test plate. The slot correlating equation was adapted to fit these data by the definition of an effective slot height. An additional corrective term was then determined to correlate these data.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-299 , E-689
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  • 25
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    Aerodynamic Characteristics of a Large-Scale Unswept Wing-Body-Tail Configuration with Blowing Applied Over the Flap and Wind Leading Edge (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: An investigation has been conducted in the Langley full-scale tunnel to determine the effects of a blowing boundary-layer-control lift-augmentation system on the aerodynamic characteristics of a large-scale model of a fighter-type airplane. The wing was unswept at the 70-percent- chord station, had an aspect ratio of 2.86, a taper ratio of 0.40, and 4-percent-thick biconvex airfoil sections parallel to the plane of symmetry. The tests were conducted over a range of angles of attack from approximately -4 deg to 23 deg for a Reynolds number of approximately 5.2 x 10(exp 6) which corresponds to a Mach number of 0.08. Blowing rates were normally restricted to values just sufficient to control air-flow separation. The results of this investigation showed that wing leading-edge blowing in combination with large values of wing leading-edge-flap deflection was a very effective leading-edge flow-control device for wings having highly loaded trailing-edge flaps. With leading-edge blowing there was no hysteresis of the lift, drag, and pitching-moment characteristics upon recovery from stall. End plates were found to improve the lift and drag characteristics of the test configuration in the moderate angle-of-attack range, and blockage to one-quarter of the blowing-slot area was not detrimental to the aerodynamic characteristics. Blowing boundary-layer control resulted in a considerably reduced landing speed and reduced landing and take-off distances. The ailerons were very effective lateral-control devices when used with blowing flaps.
    Keywords: Aerodynamics
    Type: NASA-TN-D-407 , L-927
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  • 26
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    Some Divergence Characteristics of Low-Aspect-Ratio Wings at Transonic and Supersonic Speeds (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The problem of chordwise, or camber, divergence at transonic and supersonic speeds is treated with primary emphasis on slender delta wings having a cantilever support at the trailing edge. Experimental and analytical results are presented for four wing models having apex half-angles of 5 deg, 10 deg, 15 deg, and 20 deg. A Mach number range from 0.8 to 7.3 is covered. The analytical results include calculations based on small-aspect-ratio theory, lifting-surface theory, and strip theory. A closed-form solution of the equilibrium equation is given, which is based on low-aspect-ratio theory but which applies only to certain stiffness distributions. Also presented is an iterative procedure for use with other aerodynamic theories and with arbitrary stiffness distribution.
    Keywords: Aerodynamics
    Type: NASA-TN-D-461 , L-582
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  • 27
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    Measurements of the Time-Averaged and Instantaneous Induced Velocities in the Wake of a Helicopter Rotor Hovering at High Tip Speeds (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Measurements of the time-averaged induced velocities were obtained for rotor tip speeds as great as 1,100 feet per second (tip Mach number of 0.98) and measurements of the instantaneous induced velocities were obtained for rotor tip speeds as great as 900 feet per second. The results indicate that the small effects on the wake with increasing Mach number are primarily due to the changes in rotor-load distribution resulting from changes in Mach number rather than to compressibility effects on the wake itself. No effect of tip Mach number on the instantaneous velocities was observed. Under conditions for which the blade tip was operated at negative pitch angles, an erratic circulatory flow was observed.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-393 , L-836
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  • 28
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    Free-Spinning-Tunnel Investigation of a 1/20-Scale Model of an Unswept-Wing Jet-Propelled Trainer Airplane (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A flutter analysis employing the kernel function for three- dimensional, subsonic, compressible flow is applied to a flutter-tested tail surface which has an aspect ratio of 3.5, a taper ratio of 0.15, and a leading-edge sweep of 30 deg. Theoretical and experimental results are compared at Mach numbers from 0.75 to 0.98. Good agreement between theoretical and experimental flutter dynamic pressures and frequencies is achieved at Mach numbers to 0.92. At Mach numbers from 0.92 to 0.98, however, a second solution to the flutter determinant results in a spurious theoretical flutter boundary which is at a much lower dynamic pressure and at a much higher frequency than the experimental boundary.
    Keywords: Aerodynamics
    Type: NASA-TN-D-381 , L-872
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  • 29
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    Preliminary Full-Scale Power-Off Drag of the X-15 Airplane for Mach Numbers from 0.7 to 3.1 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Drag characteristics have been obtained for the X-15 airplane during unpowered flight. These data represent a Mach number range from about 0.7 to 3.1 and a Reynolds number range from 13.9 x 10(exp 6) to 28 x 10(exp 8), based on the mean aerodynamic chord. The full-scale data are compared with estimates compiled from several wind-tunnel facilities. The agreement between wind-tunnel and full-scale supersonic drag, uncorrected for Reynolds number effects, is reasonably close except at low supersonic Mach numbers where the flight values are significantly higher.
    Keywords: Aerodynamics
    Type: NASA-TM-X-430
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  • 30
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    Investigation at Mach Numbers of 0.20 to 3.50 of Blended Wing-Body Combinations of Sonic Design with Diamond, Delta, and Arrow Plan Forms (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The models had aspect-ratio-2 diamond, delta, and arrow wings with the leading edges swept 45.00 deg, 59.04 deg, and 70.82 deg, respectively. The wing sections were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local chords in a streamwise direction. The models were tested with transition fixed at Reynolds numbers of about 4,000,000 to 9,000,0000, based on the mean aerodynamic chord of the wings. The effect of varying Reynolds number was checked at both subsonic and supersonic speeds. The diamond model was superior to the other plan forms at transonic speeds ((L/D)max = 11.00 to 9.52) because of its higher lift-curve slope and near optimum wave drag due to the blending process. For the wing thickness tested with the diamond model, the marked body and wing contouring required for transonic conditions resulted in a large wave-drag penalty at the higher supersonic Mach numbers where the leading and trailing edges of the wing were supersonic. Because of the low sweep of the trailing edge of the delta model, this configuration was less adaptable to the blending process. Removing a body bump prescribed by the Mach number 1.00 design resulted in a good supersonic design. This delta model with 10 percent less volume was superior to the other plan forms at Mach numbers of 1.55 to 2.35 ((L/D)max = 8.65 to 7.24), but it and the arrow model were equally good at Mach numbers of 2.50 to 3.50 ((L/D)max - 6.85 to O.39). At transonic speeds the arrow model was inferior because of the reduced lift-curve slope associated with its increased sweep and also because of the wing base drag. The wing base-drag coefficients of the arrow model based on the wing planform area decreased from a peak value of 0.0029 at Mach number 1.55 to 0.0003 at Mach number 3.50. Linear supersonic theory was satisfactory for predicting the aerodynamic trends at Mach numbers from 1.55 to 3.50 of lift-curve slope, wave drag, drag due to lift, aerodynamic-center location, and maximum lift-drag ratios for each of the models.
    Keywords: Aerodynamics
    Type: NASA-TM-X-372
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  • 31
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    Empirical Equation for Turbulent Forced-Convection Heat Transfer for Prandtl Numbers from 0.001 to 1000 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A review is made of some of the experimental data and analyses applicable to convective heat transfer in fully turbulent flow in smooth tubes with liquid metals and viscous Newtonian fluids. An empirical equation is evolved that closely approximates heat-transfer values obtained from selected analyses and experimental data for Prandtl numbers from 0.001 to 1000. The terms included in the equation are Reynolds number, Prandtl number, and an empirical diffusivity ratio between heat and momentum.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-483
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  • 32
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    NASA Research on the Hydrodynamics of the Gaseous Vortex Reactor (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The experimental and analytical results to date of a study of a two-component gaseous vortex system are presented in this paper. Analytical expressions for tangential velocity and static-pressure profiles in a turbulent vortex show good agreement with experimental data. Airflow rates from 0.075 to 0.14 pound per second and corresponding tangential velocities from 160 to 440 feet per second are correlated by turbulent Reynolds numbers from 1.95 to 2.4. An analysis of an air-bromine gas mixture in a turbulent vortex indicates that a boundary value of bromine-to-air radial velocity ratio (u(2)/u(1)) of 0.999 gives essentially no bromine buildup, while a value of 0.833 results in considerable separation. For a constant value of (u(2)/u(1))(0) the bromine buildup increases as (1) the tangential velocity increases, (2) the air-to-bromine weight-flow ratio decreases, (3) the airflow rate decreases, (4) the temperature decreases, and (5) the turbulence decreases. Analytical temperature, pressure, and tangential-velocity profiles are also presented. Preliminary experimental results indicate that the flow of an air-bromine mixture through a vortex field results in a bromine density increase to a maximum value; followed by a decrease; the air density exhibits a uniform decrease from the outer vortex radius to the exhaust-nozzle radius.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-288 , E-800 , Nov 16, 1959 - Nov 21, 1959; Washington, DC; United States
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  • 33
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    Mixing Characteristics Downstream of Core Region of High-Temperature Axisymmetric Jets Exhausting Into Transonic and Supersonic Streams (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A series of rocket motors with varying exit to throat area ratios was tested in the 8- by 6-foot wind tunnel to determine the effects of mixing on jet diameter and temperature decay at large distances (x/d 〉 30) from the nozzle exit. An approximate method to account for effects of the initial expansion was evolved. It was determined that the combustion efficiency has an important effect on jet spreading, since the unburned products can burn downstream of the nozzle. The data showed considerable scatter; however, mixing rates were, in general, lower than those observed for subsonic jets. Data for angles of attack of 5 and 10 deg are also presented, giving the respective centerline shift and temperature decay as a function of axial distance.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TM-X-151
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  • 34
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    Friction and Pressure Drag of Boundary-Layer Diverter Systems at Mach Number of 3.0 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An experimental investigation was performed at a Mach number of 3.0 to determine the friction and pressure drags of a pylon and a 20 deg- and a 40 deg-included-angle wedge diverter over a range of Reynolds number. The results indicated that the measured friction drag coefficients agreed reasonably with that predicted by flat-plate theory. The pressure drag coefficients of the 20 and 40 deg wedges agreed with those presented in the literature. The total drag coefficient of the pylon and the 20 deg wedge diverter was about 0.36, based on diverter frontal area, while the drag coefficient of the 40 deg wedge was about 0.47.
    Keywords: Aerodynamics
    Type: NASA-TM-X-147
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  • 35
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    Free-spinning-tunnel Investigation of a 1/30 Scale Model of a Twin-jet-swept-wing Fighter Airplane (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation has been made in the Langley 20-foot free-spinning tunnel to determine the erect and inverted spin and recovery characteristics of a 1/30-scale dynamic model of a twin-jet swept-wing fighter airplane. The model results indicate that the optimum erect spin recovery technique determined (simultaneous rudder reversal to full against the spin and aileron deflection to full with the spin) will provide satisfactory recovery from steep-type spins obtained on the airplane. It is considered that the air-plane will not readily enter flat-type spins, also indicated as possible by the model tests, but developed-spin conditions should be avoided in as much as the optimum recovery procedure may not provide satisfactory recovery if the airplane encounters a flat-type developed spin. Satisfactory recovery from inverted spins will be obtained on the airplane by neutralization of all controls. A 30-foot- diameter (laid-out-flat) stable tail parachute having a drag coefficient of 0.67 and a towline length of 27.5 feet will be satisfactory for emergency spin recovery.
    Keywords: Aerodynamics
    Type: NASA-TM-SX-446 , L-1191 , N5154
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  • 36
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    Ground Measurements of Airplane Shock-Wave Noise at Mach Numbers to 2.0 and at Altitudes to 60,000 Feet (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-14
    Description: The intensity of shock-wave noise at the ground resulting from flights at Mach numbers to 2.0 and altitudes to 60,000 feet was measured. Meagurements near the ground track for flights of a supersonic fighter and one flight of a supersonic bomber are presented. Level cruising flight at an altitude of 60,000 feet and a Mach number of 2.0 produced sonic booms which were considered to be tolerable, and it is reasonable t o expect that cruising flight at higher altitudes will produce booms of tolerable intensity for airplanes of the size and weight of the test airplanes. The measured variation of sonic-boom intensity with altitude was in good agreement with the variation calculated by an equation given in NASA Technical Note D-48. The effect of Mach number on the ground overpressure is small between Mach numbers of 1.4 and 2.0, a result in agreement with the theory. No amplification of the shock-wave overpressures due to refraction effects was apparent near the cutoff Mach number. A method for estimating the effect of fligh-path angle on cutoff Mach number is shown. Experimental results indicate agreement with the method, since a climb maneuver produced booms of a much decreased intensity as compared with the intensity of those measured in level flight at about the same altitude and Mach number. Comparison of sound pressure levels for the fighter and bomber airp lanes indicated little effect of either airplane size or weight at an altitude of 40,000 feet.
    Keywords: Aerodynamics
    Type: NASA-TN-D-235
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  • 37
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    Shock Losses in Transonic Compressor Blade Rows (1960)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The rather extensive study of the shock losses in transonic compressors can he summarized by the following remarks: 1. A simple flow model can be used to estimate shock losses at the design point for transonic compressor blade rows and results iii reasonable correlation of loss data. It is indicated that shock losses can constitute a sizable portion of the total losses in it transonic compressor rotor. This includes all blade elements at which sonic or higher relative velocities are obtained. 2. Shock losses can he shown to exist across the blade passage (free-stream loss) and by the method of superposition with the blade profile losses result in an estimated design total loss coefficient. 3. The shock configuration was experimentally determined by the rapid pressure rise between the blades as measured by the use of barium titanate crystals. At the minimum loss operating conditions the shock is very similar to that assumed in the simple How model. 4. Shock losses obtained from a more detailed flow model were compared with the losses obtained by the simple flow model. Measured loss distribution from blade to blade closely approaches the analytical shock loss distribution. The measured distribution shows the effect of a shock boundary layer interaction. 5. The analytical method (from the detailed flow model) of determining the shock location ahead of the blade seems to apply reasonably well over a range of incidence angles. The analytical shock losses do not vary a great deal with blade element incidence angles.
    Keywords: Aerodynamics
    Type: ASME Paper No. 60-WA-77 , ASME Winter Annual Meeting; Nov 27, 1960 - Dec 02, 1960; New York, NY; United States
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  • 38
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    High Speed Schlieren Studies of Flow Over Mercury Atlas Vehicle in the Langley 2-Foot Transonic Aeroplasticity Tunnel (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: Test conditions for the studies are: Mach number varying continuously from approximately 0.8 to 1.1 and Reynolds number (based on maximum diameter of Atlas) approximately 0.451 x 10(exp 6). Camera speed is 2000 frames per second.
    Keywords: Aerodynamics
    Type: L-583
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  • 39
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    Flow Over Blunt Body at M equals 20 in 2-Inch Helium Tunnel (1960)
    In:  Other Sources
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    Publication Date: 2019-07-12
    Description: The film shows flow over blunt body alone, with internal spike, and with external spikes.
    Keywords: Aerodynamics
    Type: L-562
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  • 40
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    High-Speed Schlieren Movies of Decelerators at Supersonic Speeds (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: Tests were conducted on several types of porous parachutes, a paraglider, and a simulated retrorocket. Mach numbers ranged from 1.8-3.0, porosity from 20-80 percent, and camera speeds from 1680-3000 feet per second (fps) in trials with porous parachutes. Trials of reefed parachutes were conducted at Mach number 2.0 and reefing of 12-33 percent at camera speeds of 600 fps. A flexible parachute with an inflatable ring in the periphery of the canopy was tested at Reynolds number 750,000 per foot, Mach number 2.85, porosity of 28 percent, and camera speed of 36oo fps. A vortex-ring parachute was tested at Mach number 2.2 and camera speed of 3000 fps. The paraglider, with a sweepback of 45 degrees at an angle of attack of 45 degrees was tested at Mach number 2.65, drag coefficient of 0.200, and lift coefficient of 0.278 at a camera speed of 600 fps. A cold air jet exhausting upstream from the center of a bluff body was used to simulate a retrorocket. The free-stream Mach number was 2.0, free-stream dynamic pressure was 620 lb/sq ft, jet-exit static pressure ratio was 10.9, and camera speed was 600 fps.
    Keywords: Aerodynamics
    Type: L-569
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  • 41
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    High-Speed Schlieren Movies of the Flow About Reefed Parachute Models Towed at Supersonic Speeds Behind a Conical Body (4.875 Inches in Diameter). Drag Values Based on the Unreefed Diameter of 1.73 F. Porosity of Unreefed Parachute is 28 Percent. (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: Flexible parachute models reefed to one-eighth, one-fourth, one-third, and four tenths of its diameter were towed at speeds of Mach 1.80, 2.00, 2.20 and 2.87. Towline lengths tested were 23.40, 24.38, 26.81, and 29.25 inches. High-speed Schlieren movies of the flow are shown.
    Keywords: Aerodynamics
    Type: L-556
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  • 42
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    Calculation of Aerodynamic Loading and Twist Characteristics of a Flexible Wing at Mach Numbers Approaching 1.0 and Comparison with Experiment (1960)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: An iteration method is presented by which the detailed aerodynamic loading and twist characteristics of a flexible wing with known elastic properties may be calculated. The method is applicable at Mach numbers approaching 1.0 as well as at subsonic Mach numbers. Calculations were made for a wing-body combination; the wing was swept back 45 deg and had an aspect ratio of 4. Comparisons were made with experimental results at Mach numbers from.0.80 to 0.98.
    Keywords: Aerodynamics
    Type: NASA-TR-R-58
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  • 43
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    Recent Hypersonic Studies of Wings and Bodies (1960)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Even though a great deal of theoretical and experimental information has been obtained in recent years on the flow over simple shapes in hypersonic flow a great deal of confusion still exists on how to interpret and extrapolate the results obtained. This paper offers information recently obtained at Langley at Mach numbers ranging from 7 to 21 encompassing both work in air and helium on shapes ranging from rods to delta wings. The results indicate that in most cases methods for making useful estimates of pressure are in hand for simple shapes. However, three-dimensional effects and the interaction between the components considerably complicates the flow fields over delta wings at low angles of attack.
    Keywords: Aerodynamics
    Type: ARS Semi-Annual Meeting; May 09, 1960 - May 12, 1960; Los Angeles, CA; United States
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  • 44
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    Effects of Fineness Ratio and Reynolds Number on the Low-Speed Crosswind Drag Characteristics of Circular and Modified-Square Cylinders (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A wind-tunnel investigation has been made on modified-square and circular cylinders to determine the effects of fineness ratio and Reynolds numbers on the crosswind drag characteristics. Fineness ratios from 2 to 14 were investigated over a Reynolds number range from approximately 300,000 to 1,650,000 which corresponded to Mach numbers from 0.057 to 0.377.The result of the investigation show that at supercraft Reynolds numbers the drag coefficient of the circular cylinder increases with increasing Reynolds number for all fineness ratios but at low fineness ratios this effect is considerably less than at higher fineness ratios. For circular cylinders in the high fineness-ratio range there is a reduction in drag as the fineness ratio is decreased except for Reynolds numbers of 900,000 and 1,000,000, whereas at low fineness ratios the opposite trend generally occurs. The addition of hemispherical ends to the circular cylinder gave a substantial decrease in drag at a fineness ratio of 3.27 but the effect was negligible at fineness ratios of 5.27 and 10. The finite-length modified-square cylinder gave the reduction in drag over the two-dimensional modified-square cylinder for the complete range of test Reynolds numbers with the lowest fineness ratio giving the lowest drag at Reynolds numbers above 3O0,OOO.
    Keywords: Aerodynamics
    Type: NASA-TN-D-540 , L-1020
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  • 45
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    A Supersonic Area Rule and an Application to the Design of a Wing-Body Combination with High Lift-Drag Ratios (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A concept for interrelating the wave drags of wing-body combinations at supersonic speeds with axial developments of cross-sectional area is presented. A swept-wing-indented-body combination designed on the basis of this concept to have significantly improved maximum lift-drag ratios over a range of transonic and moderate supersonic speeds is described. Experimental results have been obtained for this configuration at Mach numbers from 0.80 to 2.01. Maximum lift-drag ratios of approximately 14 and 9 were measured at Mach numbers of 1.15 and 1.41, respectively.
    Keywords: Aerodynamics
    Type: NASA-TR-R-72
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  • 46
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    The Numerical Calculation of Flow Past Conical Bodies Supporting Elliptic Conical Shock Waves at Finite Angles of Incidence (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The inverse method, with the shock wave prescribed to be an elliptic cone at a finite angle of incidence, is applied to calculate numerically the supersonic perfect-gas flow past conical bodies not having axial symmetry. Two formulations of the problem are employed, one using a pair of stream functions and the other involving entropy and components of velocity. A number of solutions are presented, illustrating the numerical methods employed, and showing the effects of moderate variation of the initial parameters.
    Keywords: Aerodynamics
    Type: NASA-TN-D-340 , A-385
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  • 47
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    An Experimental Study of the Ionization of Low-Density Gas Flows by Induced Discharges (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Induced discharges are advantageous for ionizing low-density flows in that they introduce no electrode contamination into the flow and they provide a relatively high degree of ionization with good coupling of power into the gas. In this investigation a 40-megacycle oscillator was used to produce and maintain induced discharges in argon and mercury-vapor flows. Methods for preventing blowout of the discharge were determined, and power measurements were made with an in-line wattmeter. Some results with damped oscillations pulsed at 1,000 pulses per second are also presented.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-431 , L-986
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  • 48
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    Equations for the Induced Velocities Near a Lifting Rotor with Nonuniform Azimuthwise Vorticity Distribution (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Equations, which can be integrated on high-speed computing machines, are developed for all three components of induced velocity at an arbitrary point near the rotor and for an arbitrary harmonic variation of vorticity. Sample calculations for vorticity which varies as the sine of the azimuth angle indicate that the normal component of induced velocity is, in this case, uniform along either side of the lateral axis.
    Keywords: Aerodynamics
    Type: NASA-TN-D-394 , L-797
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  • 49
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    An Experimental Investigation of Boundary-Layer Control for Drag Reduction of a Swept-Wing Section at Low Speed and High Reynolds Numbers (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation of laminar boundary-layer control by suction for purposes of drag reduction at low speed and high Reynolds numbers has been conducted in the Ames 12-Foot Pressure Wind Tunnel. The model was a 72.96-inch-chord wing panel, swept back 30 deg., which was installed between end plates to approximate a wing of infinite span. The airfoil section employed was a modified NACA 66-012 in the streamwise direction. Tests were limited to controlling the flow over only the upper surface of the model. Seventeen individually controllable suction chambers were provided below the surface to induce flow through 93 spanwise slots in the surface between the 0.0052- and 0.97-chord stations. Tests were made at angles of attack of 0 deg., +/- 1.0 deg., +/- 1.5 deg., and -2.0 deg. for Reynolds numbers from approximately 1.5 x 10(exp 6) to 4.0 x 10(exp 6) per foot. In general, essentially full-chord laminar flow was obtained for all conditions with small suction quantities. Minimum profile-drag coefficients of about 0.0005 to 0.0006 were obtained for the slotted surface at maximum values of the Reynolds number; these values include the Power required to induce suction as an equivalent drag.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-320
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  • 50
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    Low-Speed Aerodynamic Characteristics of a Model of a Hypersonic Research Airplane at Angles of Attack up to 90 deg for a Range of Reynolds Numbers (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Static force tests have been made at low subsonic speeds for a model of a hypersonic research airplane in the Langley high-speed 7- by 10-foot tunnel to determine the aerodynamic forces and moments up to an angle of attack of 90 deg for a range of Reynolds numbers. The Reynolds numbers, based on the mean aerodynamic chord, ranged from 740,000 to 1,900,000, which correspond to dynamic pressures from 15 to 100 lb/sq ft (Mach numbers from 0.10 to 0.27). The model was tested in the clean configuration with various horizontal-tail settings, horizontal tail off, lower rudder off, fuselage alone, and with various size strakes and slats on the nose of the model. Representative results of the present investigation are presented in plotted form, and a tabulation of all the data obtained is presented in a table. Appreciable effects on side force, yawing moment, and pitching moment are indicated by changes in Reynolds number for angles of attack of 40 to 90 deg.
    Keywords: Aerodynamics
    Type: NASA-TN-D-403 , L-905
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  • 51
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    Supersonic Aerodynamic Characteristics of a Low-Drag Aircraft Configuration having an Arrow Wing of Aspect Ratio 1.86 and a Body of Fineness Ratio 20 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A free-flight rocket-propelled-model investigation was conducted at Mach numbers of 1.2 to 1.9 to determine the longitudinal and lateral aero-dynamic characteristics of a low-drag aircraft configuration. The model consisted of an aspect-ratio -1.86 arrow wing with 67.5 deg. leading-edge sweep and NACA 65A004 airfoil section and a triangular vertical tail with 60 deg. sweep and NACA 65A003 section in combination with a body of fineness ratio 20. Aerodynamic data in pitch, yaw, and roll were obtained from transient motions induced by small pulse rockets firing at intervals in the pitch and yaw directions. From the results of this brief aerodynamic investigation, it is observed that very slender body shapes can provide increased volumetric capacity with little or no increase in zero-lift drag and that body fineness ratios of the order of 20 should be considered in the design of long-range supersonic aircraft. The zero-lift drag and the drag-due-to-lift parameter of the test configuration varied linearly with Mach number. The maximum lift-drag ratio was 7.0 at a Mach number of 1.25 and decreased slightly to a value of 6.6 at a Mach number of 1.81. The optimum lift coefficient, normal-force-curve slope, lateral-force-curve slope, static stability in pitch and yaw, time to damp to one-half amplitude in pitch and yaw, the sum of the rotary damping derivatives in pitch and also in yaw, and the static rolling derivatives all decreased with an increase in Mach number. Values of certain rolling derivatives were obtained by application of the least-squares method to the differential equation of rolling motion. A comparison of the experimental and calculated total rolling-moment-coefficient variation during transient oscillations of the model indicated good agreement when the damping-in-roll contribution was included with the static rolling-moment terms.
    Keywords: Aerodynamics
    Type: NASA-TN-D-509 , L-894
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  • 52
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    Effect of Reynolds Number on the Force and Pressure Distribution Characteristics of a Two-Dimensional Lifting Circular Cylinder (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A two-dimensional lifting circular cylinder has been tested over a Mach number range from 0.011 to 0.32 and a Reynolds number range from 135,000 to 1,580,000 to determine the force and pressure distribution characteristics. Two flaps having chords of 0.37 and 6 percent of the cylinder diameter, respectively, and attached normal to the surface were used to generate lift. A third configuration which had 6-percent flaps 1800 apart was also investigated. All flaps were tested through a range of angular positions. The investigation also included tests of a plain cylinder without flaps. The lift coefficient showed a wide variation with Reynolds number for the 6-percent flap mounted on the bottom surface at the 50-percent-diameter station, varying from a low of about 0.2 at a Reynolds number of 165,000 to a high of 1.54 at a Reynolds number of 350,000 and then decreasing almost linearly to a value of 1.0 at a Reynolds number of 1,580,000. The pressure distribution showed that the loss of lift with Reynolds number above the critical was the result of the separation point moving forward on the upper surface. Pressure distributions on a plain cylinder also showed similar trends with respect to the separation point. The variation of drag coefficient with Reynolds number was in direct contrast to the lift coefficient with the minimum drag coefficient of 0.6 occurring at a Reynolds number of 360,000. At this point the lift-drag ratios were a maximum at a value of 2.54. Tests of a flap with a chord of 0.0037 diameter gave a lift coefficient of 0.85 at a Reynolds number of 520,000 with the same lift-drag ratio as the larger flap but the position of the flap for maximum lift was considerably farther forward than on the larger flap. Tests of two 6-percent flaps spaced 180 deg apart showed a change in the sign of the lift developed for angular positions of the flap greater than 132 deg at subcriti- cal Reynolds numbers. These results may find use in application to air- craft using forebody strakes. The drag coefficient developed by the flaps when normal to the relative airstream was approximately equal to that developed by a flat plate in a similar attitude.
    Keywords: Aerodynamics
    Type: NASA-TN-D-455 , L-936
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  • 53
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    Theoretical Calculations of the Pressures, Forces, and Moments Due to Various Lateral Motions Acting on Tapered Sweptback Vertical Tails with Supersonic Leading and Trailing Edges (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Based on expressions for the linearized velocity potentials and pressure distributions given in NACA Technical Report 1268, formulas for the span load distribution, forces, and moments are derived for families of thin isolated vertical tails with arbitrary aspect ratio, taper ratio, and sweepback performing the motions constant sideslip, steady rolling, steady yawing, and constant lateral acceleration. The range of Mach number considered corresponds, in general, to the condition that the tail leading and trailing edges are supersonic. To supplement the analytical results, design-type charts are presented which enable rapid estimation of the forces and moments (expressed as stability derivatives) for given combinations of geometry parameters and Mach number.
    Keywords: Aerodynamics
    Type: NASA-TN-D-383 , L-780
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  • 54
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    Real-Gas Correction Factors for Hypersonic Flow Parameters in Helium (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The real-gas hypersonic flow parameters for helium have been calculated for stagnation temperatures from 0 F to 600 F and stagnation pressures up to 6,000 pounds per square inch absolute. The results of these calculations are presented in the form of simple correction factors which must be applied to the tabulated ideal-gas parameters. It has been shown that the deviations from the ideal-gas law which exist at high pressures may cause a corresponding significant error in the hypersonic flow parameters when calculated as an ideal gas. For example the ratio of the free-stream static to stagnation pressure as calculated from the thermodynamic properties of helium for a stagnation temperature of 80 F and pressure of 4,000 pounds per square inch absolute was found to be approximately 13 percent greater than that determined from the ideal-gas tabulation with a specific heat ratio of 5/3.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-462 , L-1135
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  • 55
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    Experimental Smoke and Electromagnetic Analog Study of Induced Flow Field About a Model Rotor in Steady Flight Within Ground Effect (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Hovering and steady low-speed forward-flight tests were run on a 4-foot-diameter rotor at a ground height of 1 rotor radius. The two blades had a 2 to 1 taper ratio and were mounted in a see-saw hub. The solidity ratio was 0.05. Measurements were made of the rotor rpm, collective pitch, and forward-flight velocity. Smoke was introduced into the tip vortex and the resulting vortex pattern was photographed from two positions. Using the data obtained from these photographs, wire models of the tip vortex configurations were constructed and the distribution of the normal component of induced velocity at the blade feathering axis that is associated with these tip vortex configurations was experimentally determined at 450 increments in azimuth position from this electromagnetic analog. Three steady-state conditions were analyzed. The first was hovering flight; the second, a flight velocity just under the wake "tuck under" speed; and the third, a flight velocity just above this speed. These corresponded to advance ratios of 0, 0.022, and 0.030 (or ratios of forward velocity to calculated hovering induced velocity of approximately 0, 0.48, and 0.65), respectively, for the model test rotor. Cross sections of the wake at 450 intervals in azimuth angle as determined from the path of the tip vortex are presented graphically for all three cases. The nondimensional normal component of the induced velocity that is associated with the tip vortex as determined by an electromagnetic analog at 450 increments in azimuth position and at the blade feathering axis is presented graphically. It is shown that the mean value of this component of the induced velocity is appreciably less after tuck-under than before. It is concluded that this method yields results of engineering accuracy and is a very useful means of studying vortex fields.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-458 , W-143
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  • 56
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    Distribution of Time-Averaged Pressure Fluctuations Along the Boundary of a Round Subsonic Jet (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A semiempirical analysis of the equation for incompressible fluctuations in a turbulent fluid, using similarity relations for round subsonic jets with uniform exit velocity, is used to predict the shape of the time-averaged fluctuation-pressure distribution along the mean-velocity boundary of jets. The predicted distribution is independent of distance downstream of the nozzle exit along the mixing region, inversely proportional to the distance downstream along the region of mean-velocity self-preservation, and proportional to the inverse square of the distance downstream along the fully developed region. Experimental results were in fair agreement with the theory. However, the measured fluctuation-pressure distributions were found to be very sensitive to changes in jet temperature and jet-nozzle profile, especially near the nozzle. These factors are not included in the theory. Increased jet temperatures produce increased pressure fluctuations and violation of similarity conditions. Nozzle-profile modifications may lead to violation of the uniform-exit-velocity requirement imposed in the theory.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-468 , E-780
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  • 57
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    The Blunt Plate In Hypersonic Flow (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The sonic-wedge characteristics method has been used to obtain the shock shapes and surface pressure distributions on several blunt two-dimensional shapes in a hypersonic stream for several values of the ratio of specific heats. These shapes include the blunt slab at angle of attack and power profiles of the form yb = a)P, where 0 les than m less than 1, Yb and x are coordinates of the body surface, and a is a constant. These numerical results have been compared with the results of blast-wave theory, and methods of predicting the pressure distributions and shock shapes are proposed in each case. The effects of a free-stream conical-flow gradient on the pressure distribution on a blunt slab in hypersonic flow were investigated by the sonic-wedge characteristics method and were found to be sizable in many cases. Procedures which are satisfactory for reducing pressure data obtained in conical flows with small gradients are presented.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-408 , L-897
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  • 58
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    Investigation of the Flow Over Simple Bodies at Mach Numbers of the Order of 20 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: It is shown that adequate means are available for calculating inviscid direct and induced pressures on simple axisymmetric bodies at zero angle of attack. The extent to which viscous effects can alter these predictions is indicated. It is also shown that inviscid induced pressures can significantly affect the stability of blunt, two-dimensional flat wings at low angles of attack. However, at high angles of attack, the inviscid induced pressure effects are negligible.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-449 , L-1051
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  • 59
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    Variation in Heat Transfer During Transient Heating of a Hemisphere at a Mach Number of 2 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Convective heat-transfer tests were made on a 5-inch-diameter hemisphere to determine the variation of Stanton number with the ratio of wall temperature to total temperature. The tests were made at a nominal Mach number of 2 for stagnation temperatures of 760 deg R, 1,030 deg R, and 1,380 deg R. The model was constructed so that radiation effects and also streamwise conduction effects within the model skin were minimized. The results of the tests verified that these effects were small. Tests which were made with different masses of air inside the model to check for conduction effects to the internal air cavity showed these effects to be negligible. For laminar flow on the hemisphere, the Stanton number remained essentially constant as the ratio of wall temperature to total temperature increased. However, for fully established turbulent flow, the Stanton number at some stations decreased on the order of 50 percent as the ratio of wall temperature to total temperature increased. A theory which agreed fairly well with the trend of this decrease is shown for comparison.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-399 , L-463
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  • 60
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    Transonic and Supersonic Wind-Tunnel Tests of Wing-Body Combinations Designed for High Efficiency at a Mach Number of 1.41 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Wind-tunnel force tests of a number of wing-body combinations designed for high lift-drag ratio at a Mach number of 1.41 are reported. Five wings and six bodies were used in making up the various wing-body combinations investigated. All the wings had the same highly swept dis- continuously tapered plan form with NACA 65A-series airfoil sections 4 percent thick at the root tapering linearly to 3 percent thick at the tip. The bodies were based on the area distribution of a Sears-Haack body of revolution for minimum drag with a given length and volume. These wings and bodies were used to determine the effects of wing twist., wing twist and camber, wing leading-edge droop, a change from circular to elliptical body cross-sectional shape, and body indentation by the area-rule and streamline methods. The supersonic test Mach numbers were 1.41 and 2.01. The transonic test Mach number range was from 0.6 to 1.2. For the transition-fixed condition and at a Reynolds number of 2.7 x 10(exp 6) based on the mean aerodynamic chord, the maximum value of lift- drag ratio at a Mach number of 1.41 was 9.6 for a combination with a twisted wing and an indented body of elliptical cross section. The tests indicated that the transonic rise in minimum drag was low and did not change appreciably up to the highest test Mach number of 2.01. The lower values of lift-drag ratio obtained at a Mach number of 2.01 can be attributed to the increase of drag due to lift with Mach number.
    Keywords: Aerodynamics
    Type: NASA-TN-D-435 , L-260
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  • 61
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    Experimental Investigation of Two Low-Drag Supercavitating Hydrofoils at Speeds up to 200 Feet per Second (1960)
    In:  CASI
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    Publication Date: 2019-08-15
    Description: An experimental investigation has been made in the Langley highspeed hydrodynamics facility to determine the force and moment characteristics of two hydrofoils (one having an aspect ratio of 1 and the other having an aspect ratio of 3) designed to have improved lift-drag ratios when operating under either supercavitating or ventilated conditions. Measurements were made of lift, drag, and pitching moment over a range of angles of attack from 40 to 200 for depths of submersion varying from 0 to approximately 1 chord. The range of speed for the investigation was from 110 to 200 feet per second. When the upper surface of the hydrofoils was completely unwetted, the experimental values of lift and drag forces were in good agreement with the theoretical values obtained from the zero-cavitation-number theory. The theoretical values for minimum angle of attack for operation with the upper surface of the hydrofoil unwetted define the lower limits of angle of attack for which the experimental values of lift coefficient are either in agreement with or slightly greater than those predicted by theory.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-436 , L-913
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  • 62
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    Preliminary Heat-Transfer Measurements on a Hypersonic Glide Configuration Having 79.5 degree Sweepback and 45 degree Dihedral at a Mach Number of 4.95 (1960)
    In:  CASI
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    Publication Date: 2019-08-15
    Description: An experimental investigation was conducted to evaluate the heat-transfer characteristics of a hypersonic glide configuration having 79.5 deg of sweepback (measured in the plane of the leading edges) and 45 of dihedral. The tests were conducted at a nominal Mach number of 4.95 and a stagnation temperature of 400 F. The test-section unit Reynolds number was varied from 1.95 x 10(exp 6) to 12.24 x 10(exp 6) per foot. The results indicated that the laminar-flow heat-transfer rate to the lower surface of the model decreased as the distance from the ridge line increased except for thermocouples located near the semispan at an angle of attack of 00 with respect to the plane of the leading edges. The heat-transfer distribution (local heating rate relative to the ridge-line heating rate) was similar to the theoretical heat-transfer distribution for a two-dimensional blunt body, if the ridge line was assumed to be the stagnation line, and could be predicted by this theory provided a modified Newtonian pressure distribution was used. Except in the vicinity of the apex, the ridge-line heat-transfer rate could also be predicted from two-dimensional blunt-body heat-transfer theory provided it was assumed that the stagnation-line heat-transfer rate varied as the cosine of the effective sweep (sine of the angle of attack of the ridge line). The heat-transfer level on the lower surface and the nondimensional heat-transfer distribution around the body on the lower surface were in qualitative agreement with the results of a geometric study of highly swept delta wings with large positive dihedrals made in reference 1.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TM-X-247
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  • 63
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    Experimental and Theoretical Study of a Rectangular Wing in a Vortical Wake at Low Speed (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A systematic study has been made, experimentally and theoretically, of the effects of a vortical wake on the aerodynamic characteristics of a rectangular wing at subsonic speed. The vortex generator and wing were mounted on a reflection plane to avoid body-wing interference. Vortex position, relative to the wing, was varied both in the spanwise direction and normal to the wing. Angle of attack of the wing was varied from -40 to +60. Both chordwise and spanwise pressure distributions were obtained with the wing in uniform and vortical flow fields. Stream surveys were made to determine the flow characteristics in the vortical wake. The vortex-induced lift was calculated by several theoretical methods including strip theory, reverse-flow theory, and reverse-flow theory including a finite vortex core. In addition, the Prandtl lifting-line theory and the Weissinger theory were used to calculate the spanwise distribution of vortex-induced loads. With reverse-flow theory, predictions of the interference lift were generally good, and with Weissinger's theory the agreement between the theoretical spanwise variation of induced load and the experimental variation was good. Results of the stream survey show that the vortex generated by a lifting surface of rectangular plan form tends to trail back streamwise from the tip and does not approach the theoretical location, or centroid of circulation, given by theory. This discrepancy introduced errors in the prediction of vortex interference, especially when the vortex core passed immediately outboard of the wing tip. The wake produced by the vortex generator in these tests was not fully rolled up into a circular vortex, and so lacked symmetry in the vertical direction of the transverse plane. It was found that the direction of circulation affected the induced loads on the wing either when the wing was at angle of attack or when the vortex was some distance away from the plane of the wing.
    Keywords: Aerodynamics
    Type: NASA-TN-D-339
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  • 64
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    Flight Measurement of Wall-Pressure Fluctuations and Boundary-Layer Turbulence (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The results are presented for a flight test program using a fighter type jet aircraft flying at pressure altitudes of 10,000, 20,000, and 30,000 feet at Mach numbers from 0.3 to 0.8. Specially designed apparatus was used to measure and record the output of microphones and hot-wire anemometers mounted on the forward-fuselage section and wing of the airplane. Mean-velocity profiles in the boundary layers were obtained from total-pressure measurements. The ratio of the root-mean-square fluctuating wall pressure to the free-stream dynamic pressure is presented as a function of Reynolds number and Mach number. The longitudinal component of the turbulent-velocity fluctuations was measured, and the turbulence-intensity profiles are presented for the wing and forward-fuselage section. In general, the results are in agreement with wind-tunnel measurements which have been-reported in the literature. For example, the variation the square root of p(sup 2)/q times the square root of p(sup 2) is the root mean square of the wall-pressure fluctuation, and q is the free-stream dynamic pressure) with Reynolds number was found to be essentially constant for the forward-fuselage-section boundary layer, while variations at the wing station were probably unduly affected by the microphone diameter (5/8 in.), which was large compared with the boundary-layer thickness.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-280
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  • 65
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    Radiation Shielding of the Stagnation Region by Transpiration of an Opaque Gas (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The laminar compressible boundary layer in the two-dimensional and axisymmetric stagnation regions has been analyzed to show the effects of the injection of a radiation absorbing foreign gas on an incident radiation field, and on the enthalpy profiles across the boundary layer. Total heat transfer to the stagnation region is evaluated for numerous cases and the results are compared with the no shielding case. Required absorption properties of the foreign gas are determined and compared with properties of known gases.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-329
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  • 66
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    Aerodynamic Characteristics at Mach Number 2.05 of a Series of Highly Swept Arrow Wings Employing Various Degrees of Twist and Camber (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A series of arrow wings employing various degrees of twist and camber were tested in the Langley 4- by 4-foot supersonic pressure tunnel. Aerodynamic forces and moments in pitch were measured at a Mach number of 2.05 and at a Reynolds number of 4.4 x 10(exp 6) based on the mean aerodynamic chord. Three of the wings, having a leading-edge sweep angle of 70 deg. and an aspect ratio of 2.24, were designed to produce a minimum drag (in comparison with that produced for other wings in the family) at lift coefficients of 0. 0.08, and 0.16. A fourth and a fifth wing, having a 75 deg. swept leading edge and an aspect ratio of 1.65, were designed for lift coefficients of 0 and 0.16, respectively. A 70 deg. swept arrow wing with twist and camber designed for an optimum loading at a lift coefficient considerably less than that for maximum lift-drag ratio gave the highest lift-drag ratio of all the wings tested a value of 8.8 compared with a value of 8.1 for the corresponding wing without twist and camber. Two twisted and cambered wings designed for optimum loading at the lift coefficient for maximum lift-drag ratio gave only small increases in maximum lift-drag ratios over that obtained for the corresponding flat wings. However, in all cases, the lift-drag ratios obtained were far below the theoretical estimates.
    Keywords: Aerodynamics
    Type: NASA-TM-X-332-1 , L-876
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  • 67
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    Basic Pressure Measurements at a Mach Number of 1.43 on a Thin 45 deg Sweptback Highly Tapered Wing with Systematic Spanwise Twist Variations (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Pressure distributions are presented on four wings: an untwisted wing to serve as a reference, and wings with linear, quadratic and cubic twist variations along the span. All the twisted wings had 0deg twist at the 10-percent-semispan station and 6deg twist at the tip. The tests were made at a Mach number of 1.43 and covered an angle-of-attack range from -4deg to 20deg. The average Reynolds number based on the wing mean aerodynamic chord was 2.9 x 10(exp 6) during tests at a stagnation pressure of 1.0 atmosphere and 1.5 x 10(exp 6) during tests at a stagnation pressure of 0.5 atmosphere.
    Keywords: Aerodynamics
    Type: NASA-TN-D-528 , L-854
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  • 68
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    Spatial Characteristics of Water Spray Formed by Two Impinging Jets at Several Jet Velocities in Quiescent Air (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The spatial characteristics of a spray formed by two impinging water jets in quiescent air were studied over a range of nominal jet velocities of 30 to 74 feet per second. The total included angle between the 0.089-inch jets was 90 deg. The jet velocity, spray velocity, disappearance of the ligaments just before drop formation, mass distribution, and size and position of the largest drops were measured in a circumferential survey around the point of jet impingement. Photographic techniques were used in the evaluations. The distance from the point of jet impingement to ligament breakup into drops was about 4 inches on the spray axis and about 1.3 inches in the radial position +/-90 deg from the axis. The distance tended to increase slightly with increase in jet velocity. The spray velocity varied from about 99 to about 72 percent of the jet velocity for a change in circumferential position from the spray axis to the +/-80 deg positions. The percentages tended to increase slightly with an increase in jet velocity. Fifty percent of the mass was distributed about the spray axis in an included angle of slightly less than 40 deg. The effect of jet velocity was small. The largest observed drops (2260-micron or 0.090-in. diam.) were found on and about the spray axis. The size of the largest drops decreased for an increase in radial angular position, being about 1860 microns (0.074 in.) at the +/-90 deg positions. The largest drop sizes tended to decrease for an increase in jet velocity, although the velocity effect was small. A drop-size distribution analysis indicated a mass mean drop size equal to 54 percent of an extrapolated maximum drop size.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-301 , E-419
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  • 69
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    Aerodynamic Characteristics at Mach Numbers of 1.41 and 2.01 of a Series of Cranked Wings Ranging in Aspect Ratio from 4.00 to 1.74 in Combination with a Body (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A program has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel to determine the effects of certain wing plan-form variations on the aerodynamic characteristics of wing-body combinations at supersonic speeds. The present report deals with the results of tests of a family of cranked wing plan forms in combination with an ogive-cylinder body of revolution. Tests were made at Mach numbers of 1.41 and 2.01 at corresponding values of Reynolds number per foot of 3.0 x 10(exp 6) and 2.5 x 10(exp 6). Results of the tests indicate that the best overall characteristics were obtained with the low-aspect-ratio wings. Plan-form changes which involved decreasing the aspect ratio resulted in higher values of maximum lift-drag ratio, in addition to large increases in wing volume. Indications are that this trend would have continued to exist at aspect ratios even lower than the lowest considered in the present tests. Increases in the maximum lift-drag ratio of about 15 percent over the basic wing were achieved with practically no increase in drag. The severe longitudinal stability associated with the basic cranked wing was no longer present (within the limits of the present tests) on the wings of lower aspect ratio formed by sweeping forward the inboard portion of the trailing edge.
    Keywords: Aerodynamics
    Type: NASA-TM-X-172 , L-261
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  • 70
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    Measurement of Screen-Size Effects on Intensity, Scale, and Spectrum of Turbulence in a Free Subsonic Jet (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The results are reported of hot-wire anemometer measurements of the fluctuating longitudinal component of the turbulent velocities in the mean flow downstream of screens in an air jet. These measurements have been analyzed by well-established techniques to give the influence of tile screen mesh size on the turbulent intensity, scale, and the power-spectral-density. The results show a linear dependence of the intensity upon the screen mesh size for locations within the central core of the air jet. The spectral-density curves show that the screens redistribute the turbulent energy from the low frequencies (〈1000 cps) to the high frequencies (〉1000 cps). The effects of the screens are overwhelmed in the mixing region of the jet flow by the turbulence levels existing there. The large pressure drops occurring across the screens reduce the velocity of the jet as compared to the jet without screens by approximately one-third for the velocity and range of mesh sizes investigated and reported in this report. The turbulence scale is a linear function of distance from the nozzle exit and is somewhat greater than comparable jets without screens.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-297 , E-798
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  • 71
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    The Influence of Low Wall Temperature on Boundary-Layer Transition and Local Heat Transfer on 2-Inch-Diameter Hemispheres at a Mach Number of 4.95 and a Reynolds Number per Foot of 73.2 x 10(exp 6) (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Measurements of the location of boundary-layer transition and the local heat transfer have been made on 2-inch-diameter hemispheres in the Langley gas dynamics laboratory at a Mach number of 4.95, a Reynolds number per foot of 73.2 x 10(exp 6), and a stagnation temperature of approximately 400 F. The transient-heating thin-skin calorimeter technique was used, and the initial values of the wall-to-stream stagnation- temperature ratios were 0.16 (cold-model tests) and 0.65 (hot-model test). During two of the four cold tests, the boundary-layer flow changed from turbulent to laminar over large regions of the hemisphere as the model heated. On the basis of a detailed consideration of the magnitude of roughness possibly present during these two cold tests, it appears that this destabilizing effect of low wall temperatures (cooling) was not caused by roughness as a dominant influence. This idea of a decrease in boundary-layer stability with cooling has been previously suggested. (See, for example, NASA Memorandum 10-8-58E.) For the laminar data obtained during the early part of the hot test, the correlation of the local-heating data with laminar theory was excellent.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-391 , L-752
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  • 72
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    Effect of Afterbody Terminal Fairings on the Performance of a Pylon-Mounted Turbojet-Nacelle Model (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: An investigation of the effect of afterbody terminal fairings on the performance of a pylon-mounted turbojet-nacelle model has been conducted in the Langley 16-foot transonic tunnel. A basic afterbody having a boattail angle of 16 deg was investigated with and without terminal fairings. The equivalent boattail angle, based on the cross-sectional area of the afterbody and terminal fairings, was 8 deg. Therefore, a simple body of revolution with a boattail angle of 8 deg was included for comparison. The tests were made at an angle of attack of 0 deg, Mach numbers of 0.80 to 1.05, jet total-pressure ratio of 1 to approximately 5, and an average Reynolds number per foot of 4.1 x 10(exp 6). A hydrogen peroxide jet simulator was used to supply the hot-jet exhaust. The results indicate that addition of terminal fairings to a 16 deg boattail afterbody increased the thrust-minus-drag coefficients and provided the lowest effective drag of the three configurations tested.
    Keywords: Aerodynamics
    Type: NASA-TM-X-215
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  • 73
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    Turbulence Studies of a Rectangular Slotted Noise-Suppressor Nozzle (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The problem of noise suppression of turbojet engines has shown a need for turbulence data within the flow field of various types of nozzles used in ad hoc investigations of the sound power. The result of turbulence studies in a nozzle configuration of four parallel rectangular slots is presented in this report with special attention to the effect of the spacing of the nozzles on the intensity of turbulence, scale of turbulence, spectrum of turbulence, and the mean stream velocity. Taylor's hypothesis, which describes the convection of the turbulence eddies, was tested and found correct within experimental error and certain experimental and theoretical limitations. The convection of the pressure patterns was also investigated, and the value of the convection velocity was found to be about 0.43 times the central core velocity of the jets. The effect of the spacing-to-width ratio of the nozzles upon the turbulence intensity, the scale of turbulence, and the spectral distribution of the noise was found in general to produce a maximum change for spacing-to-width ratios of 1.5 to 2.0. These changes may be the cause of the reduction in sound power reported for similar full-scale nozzles and test conditions under actual (static) engine operation. A noise reduction parameter is defined from Lighthill's theory which gives qualitative agreement with experiments which show the noise reduction is greatest for spacing-to-width ratios of 1.5 to 2.0.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-294 , E-384
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  • 74
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    Full-Scale Wind-Tunnel Tests of Blowing Boundary-Layer Control Applied to a Helicopter Rotor (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A full-scale wind-tunnel test was conducted of two boundary-layer-control applications to a 44-foot diameter helicopter rotor. Blowing from a nozzle near the leading edge of the blades delayed retreating blade stall. Results also indicated that delay of retreating blade stall could be obtained by cyclic blowing with a lower flow rate than that required for continuous blowing. It was found that blowing applied through a nozzle at mid-chord had no effect on retreating blade stall.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-335 , A-380
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  • 75
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    Investigation at Mach Numbers of 0.60 to 3.50 of Blended Wing-Body Combinations with Cambered and Twisted Wings with Diamond, Delta and Arrow Plan Forms (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: This investigation is a continuation of the experimental and theoretical evaluation of blended wing-body combinations. The basic diamond, delta, and arrow plan forms which had an aspect ratio of 2 with leading-edge sweeps of 45.00 deg., 59.04 deg., and 70.82 deg. and trailing edge of -45.00 deg., -18.43 deg., and 41.19 deg., respectively, are used herein as standards for evaluating the effects of camber and warp. The wing thickness distributions were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing camber and warp were computed to try to obtain nearly elliptical spanwise and chordwise load distributions for each plan form and thus to obtain low drag due to lift for a range of Mach numbers for which the velocities normal to the wing leading edge are subsonic. Elliptical chordwise load distributions were not possible for the plan forms and design conditions selected, so these distributions were somewhat different for each plan form. The models were tested with transition fixed at Mach numbers from 0.60 to 3.50 and at Reynolds numbers, based on the mean aerodynamic chord of the wing, of roughly 4,000,000 to 9,000,000. At speeds where the velocities normal to the wing leading edges were supersonic, an increase in the experimental wave-drag coefficients due to camber and twist was evident, but this penalty decreased with increased sweep. Thus the minimum wave-drag coefficients for the cambered arrow model were almost identical with the zero-lift wave- drag coefficients for the uncambered arrow model at all test Mach numbers.
    Keywords: Aerodynamics
    Type: NASA-TM-X-390
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  • 76
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    Calculated Effects of Body Shape on the Bow-Shock Overpressures in the Far Field of Bodies in Supersonic Flow (1960)
    In:  CASI
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    Publication Date: 2019-08-16
    Description: A theory for the supersonic flow about bodies in uniform flight in a homogeneous medium is reviewed and an integral which expresses the effect of body shape upon the flow parameters in the far field is reduced to a form which may be readily evaluated for arbitrary body shapes. This expression is then used to investigate the effect of nose angle, fineness ratio, and location of maximum body cross section upon the far-field pressure jump across the bow-shock of slender bodies. Curves are presented showing the variation of the shock strength with each of these parameters. It is found that, for a wide variety of shapes having equal fineness ratios, the integral has nearly a constant value.
    Keywords: Aerodynamics
    Type: NASA-TR-R-76
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  • 77
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    Low-Speed Investigation of a Full-Span Internal-Flow Jet-Augmented Flap on a High-Wing Model with a 35 deg Swept Wing of Aspect Ratio 7.0 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation of a full-span 17-percent-chord internal-flow jet-augmented flap on an aspect-ratio-7.0 wing with 35 deg of sweepback has been made in the Langley 300-MPH 7- by 10-foot tunnel. Blowing over the conventional elevator and blowing down from a nose jet were investigated as a means of trimming the large diving moments at the high momentum and high lift coefficients. The results of the investigation showed that the model with the horizontal tail 0.928 mean aerodynamic chord above the wing-chord plane was stable to the maximum lift coefficient. The large diving-moment coefficients could be trimmed either with a downward blowing nose jet or by blowing over the elevator. Neither the downward blowing nose jet nor blowing over the elevator greatly affected the static longitudinal stability of the model. Trimmed lift coefficients up to 8.8 with blowing over the elevator and up to 11.4 with blowing down at the nose were obtained when the flap was deflected 70 deg and the total momentum coefficients were 3.26 and 4.69.
    Keywords: Aerodynamics
    Type: NASA-TN-D-434 , L-931
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  • 78
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    Tests of Aerodynamically Heated Multiweb Wing Structures in a Free Jet at Mach Number 2: Five Aluminum-Alloy Models of 20-Inch Chord with 0.064-Inch-Thick Skin, 0.025-Inch-Thick Webs, and Various Chordwise Stiffening at 2 deg Angle of Attack (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An experimental study was made on five 2024-T3 aluminum-alloy multiweb wing structures (MW-2-(4), MW-4-(3), mw-16, MW-17, and MW-18), at a Mach number of 2 and an angle of attack of 2 deg under simulated supersonic flight conditions. These models, of 20-inch chord and semi-span and 5-percent-thick circular-arc airfoil section, were identical except for the type and amount of chordwise stiffening. One model with no chordwise ribs between root and tip bulkhead fluttered and failed dynamically partway through its test. Another model with no chordwise ribs (and a thinner tip bulkhead) experienced a static bending type of failure while undergoing flutter. The three remaining models with one, two, or three chordwise ribs survived their tests. The test results indicate that the chordwise shear rigidity imparted to the models by the addition of even one chordwise rib precludes flutter and subsequent failure under the imposed test conditions. This paper presents temperature and strain data obtained from the tests and discusses the behavior of the models.
    Keywords: Aerodynamics
    Type: NASA-TM-X-186
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  • 79
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    The importance of quantity of nerve fibers in development of nerve-induced supernumerary limbs in Triturus and enhancement of the nervous influence by tissue implantsSupported by a grant from the U. S. Public Health Service, Department of Health, Education and Welfare, RG-5820(C1). (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 47-59 
    Details
    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    URL: http://dx.doi.org/10.1002/jmor.1051070104
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    Masthead (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960) 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/jmor.1051060101
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    The application of Bergmann's and Allen's rules to the poikilotherms (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 85-108 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    Melanotic tumor formation in drosophilaThis investigation was begun with the aid of a research grant G-3381 from the National Science Foundation, and continued with the support of research grant RG-5285 from the National Institutes of Health, Public Health Service. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 147-157 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    URL: http://dx.doi.org/10.1002/jmor.1051060203
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    Masthead (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960) 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/jmor.1051060301
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    Histology of the reproductive organs of nutria, myocastor coypus (molina)Supported in part by a grant-in-aid, RG-3775, from the Public Health Service, and partly by a General Research Grant from the Graduate Council of Oregon State College, and published under the auspices of the Oregon State College Monographs Committee Research Paper No. 379, Department of Zoology, School of Science. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 277-299 
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    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    Fluorescence microscopy of a frog renal adenocarcinomaThis work was supported in part by a grant in aid VI-28 from the American Cancer Society. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 1-23 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    URL: http://dx.doi.org/10.1002/jmor.1051070102
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    The vocal organ of the Howling monkey (Alouatta palliata)Aided by a grant from the Wenner-Gren Foundation for Anthropological Research. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 123-140 
    Details
    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    Masthead (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960) 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    Anatomy and histochemistry of the salivary glands of the vampire bat, desmodus rotundus murinusThis study was supported by research grants (G-2188, G-7474) from the National Science Foundation to Dr. W. A. Wimsatt, and by the Dr. Solon P. Sackett and Sarah Manning Sage Research Funds of the Department of Zoology.Portion of a thesis submitted to the Graduate School in partial fulfillment of the requirements for the M.A. degree. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 301-335 
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    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    The secretory cycle of iodoproteins in ammocoetes. II. A radioautographic study of the transforming larval thyroid glandPart of a thesis submitted in partial fulfillment of the requirements for the degree of Doctor of Philosophy in the Faculty of Pure Science of Columbia University, New York. This study was aided in part by a grant from the National Science Foundation. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 357-364 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    URL: http://dx.doi.org/10.1002/jmor.1051060307
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    Adaptations for amphibious vision in the dipper (Cinclus mexicanus)Based mainly on a portion of a thesis submitted in partial fulfillment of the requirements for the degree of Doctor of Philosophy, University of Washington, Seattle, Washington. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 79-91 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
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    URL: http://dx.doi.org/10.1002/jmor.1051070106
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    Addendum (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 151-151 
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    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
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    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/jmor.1051070204
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    Hypo- and hyperdevelopment of the male genital apparatus and the Bd-M combination in Drosophila melanogaster (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 227-232 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    URL: http://dx.doi.org/10.1002/jmor.1051070207
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    Studies on the fresh water fairy shrimp Chirocephalopsis bundyi (Forbes)A revision of a thesis submitted to the faculty of Cornell University in partial fulfillment for the Ph.D. degree.. II. Histochemistry of egg-shell formation (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 259-283 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
    Additional Material: 4 Ill.
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    URL: http://dx.doi.org/10.1002/jmor.1051070303
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    The ultramicroscopic structure of a reptilian kidneySupported by grants (RG 4706 and 5479) from the National Institutes of Health, United States Public Health Service. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 205-241 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    URL: http://dx.doi.org/10.1002/jmor.1051060207
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    The spermatozoa of some north american bats and rodents (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 77-83 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
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    URL: http://dx.doi.org/10.1002/jmor.1051060103
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    Studies on transplanted embryonic limbs of the chick. VI. The development of muscles and tendons in nerveless and weakly innervated chick limb graftsSupported in part by the State College of Washington Committee on Research.For the skillful preparation of the plates the authors are indebted to Mrs. Robert F. Lanzillotti. The authors also acknowledge the helpful suggestions of Dr. George E. Hudson throughout the course of the study. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 131-146 
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    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
    Type of Medium: Electronic Resource
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    Electronic Resource
    Morphology and behavior of waltzer-type mice (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 197-203 
    Details
    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
    Additional Material: 1 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/jmor.1051060206
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  • 98
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    A study of abnormalities in parthenogenetically developing germ discs and embryos of the domestic turkeyScientific Paper No. 1877. Washington Agricultural Experiment Stations, Pullman. Project Nos. 803 and 1109.This investigation was supported, in part, by funds from Biological and Medical Research by State of Washington Initiative No. 171, and by Federal funds for regional research (W-7) under the Hatch Amended Act. (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 106 (1960), S. 263-276 
    Details
    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
    Additional Material: 12 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/jmor.1051060303
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  • 99
    Electronic Resource
    Electronic Resource
    The endocrine organs of the collared lemming (Dicrostonyx) in captivity under diverse temperature and light conditions (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 25-45 
    Details
    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
    Additional Material: 4 Tab.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/jmor.1051070103
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  • 100
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    Electronic Resource
    Gametogenesis and fertilization in the frog lung fluke, Haematoloechus medioplexus Stafford (Trematoda: Plagiorchiidae) (1960)
    New York, NY : Wiley-Blackwell
    Journal of Morphology 107 (1960), S. 93-121 
    Details
    ISSN: 0362-2525
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Medicine
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/jmor.1051070107
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