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  • Aircraft Stability and Control  (17)
  • FLUID MECHANICS AND HEAT TRANSFER  (10)
  • AERODYNAMICS  (8)
  • 1950-1954  (35)
  • 1953  (25)
  • 1950  (10)
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  • 1950-1954  (35)
Year
  • 1
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    NACA Conference on Aerodynamics of High- speed Aircraft A compilation of the papers presented (1953)
    In:  CASI
    Details
    Publication Date: 2019-05-29
    Description: Conference on aerodynamics of high speed aircraft
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-57121
    Format: application/pdf
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  • 2
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    Drag measurements at low lift of a four- nacelle airplane configuration having a longitudinal distribution of cross-sectional area conductive to low transonic drag rise (1953)
    In:  CASI
    Details
    Publication Date: 2019-05-23
    Description: Drag measurements at low lift of four-nacelle aircraft configuration with longitudinal distribution of cross-sectional area conducive to low transonic drag rise
    Keywords: AERODYNAMICS
    Type: NACA-RM-L53E29
    Format: application/pdf
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  • 3
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    Flight Determination of the Static Longitudinal Stability Boundaries of the Bell X-5 Research Airplane with 59 Deg Sweepback (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: During the flight program on the Bell X-5 airplane with 59 deg sweepback to determine the practical Mach number and normal-force coefficient limits of this configuration, a reduction in static longitudinal stability was encountered in maneuvering flight. A determination of the boundary for reduction of longitudinal stability extending to a Mach number of 0.98 is presented in this paper. A reduction of static longitudinal stability existed for all elevator and all stabilizer-executed maneuvers. The reduction of stability existed for maneuvers executed with elevator near a normal-force coefficient of 0.6 for a Mach number range of about 0.31 to 0.76. Above a Mach number of 0.76 the normal-force coefficient for reduction of stability gradually decreased to a value of 0.2 at a Mach number of 0.98. For stabilizer-executed maneuvers the stability boundary was the same as for elevator maneuvers up to a Mach number of 0.88. Above this Mach number the reduction of stability occurred at slightly higher normal-force coefficients decreasing from about 0.51 at a Mach number of 0.92 to a value of 0.311 at a Mach number of 0.97. The airplane has been flown to a Mach number of 1.04 at a normal-force coefficient of about 0.15 without encountering any reduction of stability. The pilot did not consider the reduction of stability to be dangerous at altitudes above 30,000 feet; however, precise flight was impossible. At angles of attack above that at which the reduction of longitudinal stability occurred, directional instability and aileron control overbalance were encountered.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-L53A09b
    Format: application/pdf
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  • 4
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    Preliminary Results of Stability and Control Investigation of the Bell X-5 Research Airplane (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: During the acceptance tests of the Bell X-5 airplane, measurements of the static stability and control characteristics and horizontal-tail loads were obtained by the NACA High-Speed Flight Research Station. The results of the stability and control measurements are presented in this paper. A change in sweep angle between 20 deg and 59 deg had a minor effect on the longitudinal trim, with a maximum change of about 2.5 deg in elevator deflection being required at a Mach number near 0.85; however, sweeping the wings produced a total stick-force change of about 40 pounds. At low Mach numbers there was a rapid increase in stability at high normal-force coefficients for both 20 0 and 1100 sweepback, whereas a condition of neutral stability existed for 58 0 sweepback at high normal-force coefficients. At Mach numbers near 0.8 there was an instability at normal-force coefficients above 0.5 for all sweep angles tested. In the low normal-force-coefficient range a high degree of stability resulted in high stick forces which limited the maximum load factors attainable in the demonstration flights to values under 5g for all sweep angles at a Mach number near 0.8 and an altitude of 12,000 feet. The aileron effectiveness at 200 sweepback was found to be low over the Mach number range tested.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-L52K18b
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  • 5
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    Flight Measurements of the Stability Characteristics of the Bell X-5 Research Airplane in Sideslips at 59 Deg Sweepback (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Flight measurements of the stability characteristics of the Bell X-5 research airplane at 59 deg sweepback were made in steady sideslips at Mach numbers from 0.62 to 0.97 at altitudes ranging between 35,000 and 40,000 feet. The results showed that the apparent directional stability was positive and increased at Mach numbers above 0.90. The apparent effective dihedral was positive and high, increasing at Mach numbers above 0.75. The cross-wind force coefficient per degree of sideslip was positive and increased rapidly at Mach numbers above 0.94.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-L52K13b
    Format: application/pdf
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  • 6
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    A method for calculating the lift and center of pressure of wing-body-tail combinations at subsonic, transonic, and supersonic speeds (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Keywords: AERODYNAMICS
    Type: NACA-RM-A53G08
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  • 7
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    Analogy Between Mass and Heat Transfer with Turbulent Flow (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An analysis of combined heat and mass transfer from a flat plate has been made in terms of Prandtl t s simplified physical concept of the turbulent boundary layer. The results of the analysis show that for conditions of reasonably small heat and mass transfer, the ratio of the mass-and heat-transfer coefficients is dependent on the Reynolds number of the boundary layer, the Prandtl number of the medium of diffusion, and the Schmidt number of the diffusing fluid in the medium of diffusion. For the particular case of water evaporating into air, the ratio of mass-transfer coefficient to heat-transfer coefficient is found to be slightly greater than unity.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-TN-3045
    Format: application/pdf
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  • 8
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    Impingement of Water Droplets on a Cylinder in an Incompressible Flow Field and Evaluation of Rotating Multicylinder Method for Measurement of Droplet-Size Distribution, Volume-Median Droplet-Size, and Liquid-Water Content in Clouds (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-TN-2904
    Format: application/pdf
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  • 9
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    Impingement of cloud droplets on aerodynamic bodies as affected by compressibility of air flow around the body (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-TN-2903
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  • 10
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    Analytical Investigation of Icing Limit for Diamond-Shaped Airfoil in Transonic and Supersonic Flow (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Calculations have been made for the icing limit of a diamond airfoil at zero angle of attack in terms of the stream Mach number, stream temperature, and pressure altitude. The icing limit is defined as a wetted-surface temperature of 320 F and is related to the stream conditions by the method of Hardy. The results show that the point most likely to ice on the airfoil lies immediately behind the shoulder and is subject to possible icing at Mach numbers as high as 1.4.
    Keywords: AERODYNAMICS
    Type: NACA-TN-2861
    Format: application/pdf
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  • 11
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    De-icing and runback characteristics of three cyclic electric, external de-icing boots employing chordwise shedding (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Keywords: AERODYNAMICS
    Type: NACA-RM-E53C26
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  • 12
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    Effect of Ice and Frost Formations on Drag of NACA 65(sub 1) -212 Airfoil for Various Modes of Thermal Ice Protection (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The effects of primary and. runback icing and frost formations on the drag of an 8-foot-chord NACA 651-212 airfoil section were investigated over a range of angles of attack from 20 to 80 and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.25 to 1.4 grams per cubic meter and datum air temperatures of -30 to 30 F. The results showed that glaze-ice formations, either primary or runback, on the upper surface near the leading edge of the airfoil caused large and rapid increases in drag, especially at datum air temperatures approaching 32 F and in the presence of high rates of water catch. Ice formations at lower temperatures (rime ice) did not appreciably increase the drag coefficient over the initial (standard roughness) drag coefficient. Cyclic de-icing of the primary Ice formations on the airfoil leading-edge section permitted the drag coefficient to return almost to the bare airfoil drag value. Runback icing on the lower surface did not present a serious drag problem except when heavy spanwise ridges of runback ice occurred aft of the heatable area. Frost formations caused rapid and large increases in drag with incipient stalling of the airfoil.
    Keywords: AERODYNAMICS
    Type: NACA-TN-2962
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  • 13
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    Preliminary Results of Heat Transfer from a Stationary and Rotating Ellipsoidal Spinner (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Convective heat-transfer coefficients in dry air were obtained for an ellipsoidal spinner of 30-inch maximum diameter for both stationary and rotating operation over a range of conditions including airspeeds up to 275 miles per hour, rotational speeds up to 1200 rpm, and angles of attack of zero and 40 The results are presented in terms of Nusselt numbers, Reynolds numbers, and convective heat-transfer coefficients. The studies included both uniform heating densities over the spinner and uniform surface temperatures.. In general, the results showed that rotation will increase the convective heat transfer from a spinner, especially in the turbulent-flow regions. Rotation of the spinner at 1200 rpm and at a free-stream velocity of 275 miles per hour increased the Nusselt number parameter in the turbulent-flow region by 32 percent over that obtained with a stationary spinner; whereas in the nose region, where the flow was laminar, an increase of only 18 percent was observed. Transition from laminar to turbulent flow occurred over a large range of Reynolds numbers primarily because of surface roughness of the spinner. Operation at an angle of attack of 40 had only small effects on the local convective heat transfer for the model studied.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-RM-E53F02
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  • 14
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    A Method for Rapid Determination of the Icing Limit of a Body in Terms of the Stream Conditions (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The effects of existing frictional heating were analyzed to determine the conditions under which ice formations on aircraft surfaces can be prevented. A method is presented for rapidly determining by means of charts the combination of-Mach number, altitude, and stream temperature which will maintain an ice-free surface in an icing cloud. The method can be applied to both subsonic and supersonic flow. The charts presented are for Mach numbers up to 1.8 and pressure altitudes from sea level to 45,000 feet.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-TN-2914
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  • 15
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    A Method for Determining Cloud-Droplet Impingement on Swept Wings (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The general effect of wing sweep on cloud-droplet trajectories about swept wings of high aspect ratio moving at subsonic speeds is discussed. A method of computing droplet trajectories about yawed cylinders and swept wings is presented, and illustrative droplet trajectories are computed. A method of extending two-dimensional calculations of droplet impingement on nonswept wings to swept wings is presented. It is shown that the extent of impingement of cloud droplets on an airfoil surface, the total rate of collection of water, and the local rate of impingement per unit area of airfoil surface can be found for a swept wing from two-dimensional data for a nonswept wing. The impingement on a swept wing is obtained from impingement data for a nonswept airfoil section which is the same as the section in the normal plane of the swept wing by calculating all dimensionless parameters with respect to flow conditions in the normal plane of the swept wing.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-TN-2931
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  • 16
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    Impingement of Droplets in 90 deg Elbows with Potential Flow (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Trajectories were determined for droplets in air flowing through 90 deg elbows especially designed for two-dimensional potential motion with low pressure losses. The elbows were established by selecting as walls of each elbow two streamlines of the flow field produced by a complex potential function that establishes a two-dimensional flow around a 90 deg bend. An unlimited number of elbows with slightly different shapes can be established by selecting different pairs of streamlines as walls. The elbows produced by the complex potential function selected are suitable for use in aircraft air-intake ducts. The droplet impingement data derived from the trajectories are presented along with equations in such a manner that the collection efficiency, the area, the rate, and the distribution of droplet impingement can be determined for any elbow defined by any pair of streamlines within a portion of the flow field established by the complex potential function. Coordinates for some typical streamlines of the flow field and velocity components for several points along these streamlines are presented in tabular form.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-TN-2999
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  • 17
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    Impingement of Water Droplets on NACA 65A004 Airfoil and Effect of Change in Airfoil Thickness from 12 to 4 Percent at 4 deg Angle of Attack (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The trajectories of droplets in the air flowing past an NACA 65A004 a irfoil at an angle of attack of 4 deg were determined. The amount of water in droplet form impinging on the airfoil, the area of droplet impingement, and the rate of droplet impingement per unit area on the airfoil surface were calculated from the trajectories and presented to cover a large range of flight and atmospheric conditions. The effect of a change in airfoil thickness from 12 to 4 percent at 4 deg angle of attack is presented by comparing the impingement calculations for the NACA 65A004 airfoil with those for the NACA 65(sub 1)-208 and 65(sub 1)-212 airfoils. The rearward limit of impingement on the upper surface decreases as the airfoil thickness decreases. The rearward limit of impingement on the lower surface increases with a decrease in airfoil t hickness. The total water intercepted decreases as the airfoil thickness is decreased.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-TN-3047
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  • 18
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    Investigation of Aerodynamic and Icing Characteristics of a Flush Alternate Inlet Induction System Air Scoop (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation has been made in the NACA Lewis icing research tunnel to determine the aerodynamic and icing characteristics of a full-scale induction-system air-scoop assembly incorporating a flush alternate inlet. The flush inlet was located immediately downstream of the offset ram inlet and included a 180 deg reversal and a 90 deg elbow in the ducting between inlet and carburetor top deck. The model also had a preheat-air inlet. The investigation was made over a range of mass-air- flow ratios of 0 to 0.8, angles of attack of 0 and 4 deg airspeeds of 150 to 270 miles per hour, air temperatures of 0 and 25 F various liquid-water contents, and droplet sizes. The ram inlet gave good pressure recovery in both clear air and icing but rapid blockage of the top-deck screen occurred during icing. The flush alternate inlet had poor pressure recovery in both clear air and icing. The greatest decreases in the alternate-inlet pressure recovery were obtained at icing conditions of low air temperature and high liquid-water content. No serious screen icing was observed with the alternate inlet. Pressure and temperature distributions on the carburetor top deck were determined using the preheat-air supply with the preheat- and alternate-inlet doors in various positions. No screen icing occurred when the preheat-air system was operated in combination with alternate-inlet air flow.
    Keywords: AERODYNAMICS
    Type: NACA-RM-E53E07
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  • 19
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    Low-Speed Investigation of the Static Lateral Stability and Control Characteristics of a 1/6-Scale Model of the Republic XF-84H Airplane with the Propeller Operating (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation was made to determine the static lateral stability and control characteristics of a l/6-scale model of the Republic XF-84H airplane with the propeller operating. The model had a 40deg swept wing of aspect ratio 3.45 and had a thin 3-blade supersonic-type propeller. Many modifications to the basic configuration were investigated in attempts to alleviate lateral and directional trim problems which appeared to be associated with propeller slipstream rotation. Although significant benefits were realized with several modifications, none of those tested would be expected to afford satisfactory behavior for all normal flight conditions. A marked left-wing roll-off tendency was indicated at high angles of attack for the basic model configuration. Projection of only the left slat was the most effective remedy found for this problem with the propeller operating. The use of differential wing-flap deflection also appeared to offer a promising means for reducing the roll-off tendency with power on. The large sidewash over the vertical tail, associated with slip- stream rotation, severely restricted the conditions for which directional , trim could be maintained. A small triangular dorsal fin, oriented opposite to the slipstream rotation, was found very effective in reducing the adverse sidewash flow at the tail.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL53G10
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  • 20
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    Investigation of the Low-Speed Stability and Control Characteristics of a 1/10-Scale Model of the Convair YF-102 Airplane in the Langley Free-Flight Tunnel (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation of the low-speed, power-off stability and control characteristics of a 1/10-scale model of the Convair YF-102 airplane has been made in the Langley free-flight tunnel. The model was flown over a lift-coefficient range from 0.5 to the stall in its basic configuration and with several modifications involving leading-edge slats and increases in vertical-tail size. Only relatively low-altitude conditions were simulated and no attempt was made to determine the effect of freeing the controls. The longitudinal stability characteristics of the model were considered satisfactory for all conditions investigated. The lateral stability characteristics were considered satisfactory for the basic configuration over the speed range investigated except near the stall, where large values of static directional instability caused the model to be directionally divergent. The addition of leading-edge slats or an 8-percent increase in vertical-tail area increased the angle of attack at which the model became directionally divergent. The use of leading-edge slats in combination with a 40-percent increase in vertical-tail size eliminated the directional divergence and produced satisfactory stability characteristics through the stall. The longitudinal and lateral control characteristics were generally satisfactory. Although the adverse sideslip characteristics for the model were considered satisfactory over the angle-of-attack range, analysis indicates that the adverse sideslip characteristics of the airplane may be objectionable at high angles of attack.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL53L04
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  • 21
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    Wind-Tunnel Investigation at Low Speed of the Pitching Stability Derivatives of a 1/9-Scale Powered Model of the Convair XFY-1 Vertically Rising Airplane, TED No. NACA DE 373 (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An experimental investigation has been conducted in the Langley stability tunnel at low speed to determine the pitching stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient, control deflections, and propeller blade angle were investigated. The tests were made through an angle-of-attack range from about -4deg to 29deg, and the thrust coefficient range was from 0 to 0.7. In order to expedite distribution of these data, no analysis of the data has been prepared for this paper.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL53G27
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  • 22
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    Investigation of the Spinning and Tumbling Characteristics of a 1/25-Scale Model of the Lockheed XFV-1 Airplane in the Langley 20-Foot Free-Spinning Tunnel (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation has been conducted in the Langley 20-foot free-spinning tunnel on a l/23-scale model of the Lockheed XFV-1 airplane to determine the effects of control setting and movement upon the erect-spin and recovery characteristics for a range of airplane loading conditions. A windmilling propeller was simulated on the model for some of the tests. The investigation included determination of the size of tail parachute required for emergency recovery from demonstration spins. The tumbling tendencies of the model were also investigated. The results indicated that any erect or inverted spin obtained on the airplane will be satisfactorily terminated if recovery is attempted by full rudder reversal accompanied by simultaneous lateral and longitudinal movement of the stick to neutral, The model test results showed that an 11.5-foot flat-type tail parachute (drag coefficient approximately 0.73) with a 27.5-foot towline will be effective as an emergency spin-recovery device during demonstration spins of the airplane. The model results also indicate that the airplane will not tumble for any.loading condition indicated possible.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL53G24
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  • 23
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    Flight Investigation of the Stability and Control Characteristics of a 0.13-Scale Model of the Convair XFY-1 Vertically Rising Airplane During Constant-Altitude Transitions, TED No. NACA DE 368 (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation is being conducted to determine the dynamic stability and control characteristics of a 0.13-scale flying model of the Convair XFY-1 vertically rising airplane. This paper presents the results of flight tests to determine the stability and control characteristics of the model during constant-altitude slow transitions from hovering to normal unstalled forward flight. The tests indicated that the airplane can be flown through the transition range fairly easily although some difficulty will probably encountered in controlling the yawing motions at angles of attack between about 60 and 40. An increase in the size of the vertical tail will not materially improve the controllability of the yawing motions in this range of angle of attack but the use of a yaw damper will make the yawing motions easy to control throughout the entire transitional flight range. The tests also indicated that the airplane can probably be flown sideways satisfactorily at speeds up to approximately 33 knots (full scale) with the normal control system and up to approximately 37 knots (full scale) with both elevons and rudders rigged to move differentially for roll control. At sideways speeds above these values, the airplane will have a strong tendency to diverge uncontrollably in roll.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL53E18
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  • 24
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    Wind-Tunnel Investigation at Low Speed of the Rolling Stability Derivatives of a 1/9-Scale Powered Model of the Convair XFY-1 Vertically Rising Airplane, TED No. NACA DE 373 (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An experimental investigation has been conducted in the Langley stability tunnel at low speed to determine the rolling stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient were investigated for the complete model and for certain components of the model. Effects of control deflections and of propeller blade angle were investigated for the complete model. Most of the tests were made through an angle-of-attack range from about -4deg to 29deg, and the thrust coefficient range was from 0 to 0.7. In order to expedite distribution of these data, no analysis of the data has been prepared for this paper.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL53E13
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  • 25
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    Wind-Tunnel Investigation at Low Speed of the Yawing Stability Derivatives of a 1/9-Scale Powered Model of the Convair XFY-1 Vertically Rising Airplane, TED No. NACA DE 373 (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An experimental investigation has been conducted in the Langley stability tunnel at low speed to deter+nine the yawing stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient were investigated for the complete model and for certain components of the model. Effects of control deflections and of propeller blade angle were investigated for the complete model. Most of the tests were made through an angle-of-attack range from about -4deg to 29deg, and the thrust coefficient range was from 0 to 0.7. In order to expedite distribution of these data, no analysis of the data has been prepared for this.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL53D01
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  • 26
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    Stall Characteristics Obtained from Flight 10 of Northrop X-4 No. 2 Airplane (USAF No. 46-677) (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: NACA instrumentation has been installed ii the X-J4 airplanes to obtain stability and control data during the acceptance tests conducted by the Northrop Aircraft Corporation. This report presents data obtained on the stalling characteristics of the airplane in the clean and gear- down configurations. The center of gravity was located at approximately 18 percent of the mean aerodynamic chord during the tests. The results indicated that the airplane was not completely stalled when stall was gradually approached during nominally U accelerated flight but that it was completely stalled during a more abruptly approached stall in accelerated flight. The stall in accelerated flight was relatively mild, and this was attributed to the nature of the variation of lift with angle of attack for the 001-614 airfoil section, the plan form of the wing, and to the fact that the initial sideslip at the stall produced (as shown by wind-tunnel tests of a model of the airplane) a more symmetrical stall pattern.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-A50A04
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  • 27
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    Penetration of Air Jets Issuing from Circular, Square, and Elliptical Orifices Directed Perpendicularly to an Air Stream (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted to determine the penetration of air jets d.irected perpendicularlY to an air stream. Jets Issuing from circular, square, and. elliptical orifices were investigated. and. the jet penetration at a position downstream of the orifice was determined- as a function of jet density, jet velocity, air-stream d.enaity, air-stream velocity, effective jet diameter, and. orifice flow coeffIcient. The jet penetrations were determined for nearly constant values of air-stream density at three tunnel-air velocities arid for a large range of Jet velocities and. densities. The results were correlated in terms of dimensionless parameters and the penetrations of the various shapes were compared. Greater penetration was obtained. with the square orifices and the elliptical orifices having an axis ratio of 4:1 at low tunnel-air velocities and low jet pressures than for the other orifices investigated. The square orifices gave the best penetrations at the higher values of tunnel-air velocity and jet total pressure.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-TN-2019
    Format: application/pdf
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  • 28
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    Investigation of Power Requirements for Ice Prevention and Cyclical De-Icing of Inlet Guide Vanes with Internal Electric Heaters (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted to determine the electric power requirements necessary for ice protection of inlet guide vanes by continuous heating and by cyclical de-icing. Data are presented to show the effect of ambient-air temperature, liquid-water content, air velocity, heat-on period, and cycle times on the power requirements for these two methods of ice protection. The results showed that for a hypothetical engine using 28 inlet guide vanes under similar icing conditions, cyclical de-icing can provide a total power saving as high as 79 percent over that required for continuous heating. Heat-on periods in the order of 10 seconds with a cycle ratio of about 1:7 resulted in the best over-all performance with respect to total power requirements and aerodynamic losses during the heat-off period. Power requirements reported herein may be reduced by as much as 25 percent by achieving a more uniform surface-temperature distribution. A parameter in terms of engine mass flow, vane size, vane surface temperature, and the icing conditions ahead of the inlet guide vanes.was developed by which an extension of the experimental data to icing conditions and inlet guide vanes, other than those investigated was possible.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NACA-RM-E50H29
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  • 29
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    The spanwise distribution of lift for minimum induced drag of wings having a given lift and a given bending moment (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The problem of the minimum induced drag of wings having a given lift and a given span is extended to include cases in which the bending moment to be supported by the wing is also given. The theory is limited to lifting surfaces traveling at subsonic speeds. It is found that the required shape of the downwash distribution can be obtained in an elementary way which is applicable to a variety of such problems. Expressions for the minimum drag and the corresponding spanwise load distributions are also given for the case in which the lift and the bending moment about the wing root are fixed while the span is allowed to vary. The results show a 15-percent reduction of the induced drag with a 15-percent increase in span as compared with results for an elliptically loaded wing having the same total lift and bending moment.
    Keywords: AERODYNAMICS
    Type: NACA-TN-2249 , Collected Works of Robert T. Jones; p 539-556
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  • 30
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    Tests of a Northrop XSSM-A-3 Missile in the Ames 40- by 80-Foot Wind Tunnel: Stability and Control (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: This report presents the results of wind-tunnel force tests which were conducted to determine the low-speed stability and control characteristics of a full-scale Northrop XSSM-A-3 missile. Tests were made through a range of angles of attack, sideslip, and control deflection, and at various Reynolds numbers. Characteristics of the complete missile are compared with the characteristics of the missile with the landing skids extended, with the vertical tail removed, and with the fuselage alone. No analysis of the data has been made in order to make the results available as soon as possible.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SA50D05
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  • 31
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    Effect of an Autopilot Sensitive to Yawing Velocity on the Lateral Stability of the Douglas D-558-II Airplane (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A theoretical investigation has been made to determine the effect on the lateral stability of the Douglas D-58-II airplane of an autopilot sensitive to yawing velocity. The effects of inclination of the gyro spin axis to the flight path and of tire lag in the autopilot were also determined. The flight conditions investigated included landing at sea level, approach condition at 12,000 feet, and cruising at 50,000 feet at Mach numbers of 0.80 and 1.2. The results of the investigation indicated that the lateral stability characteristics of the D-558-II airplane for the flight condition discussed should satisfy the Air Force - Navy period-damping criterion when the proposed autopilot is installed. Airplane motions in sideslip subsequent to a disturbance in sideslip are presented for several representative flight conditions in which a time lag in the autopilot of 0.10 second was assumed.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-L50F22
    Format: application/pdf
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  • 32
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    Investigations of Lateral Stability of a Glide Bomb Using Automatic Control Having No Time Lag (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The investigation of the lateral stability of an automatically controlled glide bomb led also to the attempt of clarifying the influence of a phugoid oscillation or of any general longitudinal oscillation on the lateral stability of a glide bomb. Under the assumption that its period of oscillation considerably exceeds the rolling and yawing oscillation and that c(sub a) is, at least in sections, practically constant, the result of this test is quite simple. It becomes clear that the influence of the phugoid oscillation may be replaced by suitable variation of the rolling-yawing moment on a rectilinear flight path instead of the phugoid oscillation. If the flying weight of the glide bomb of unchanged dimensions is increased, an increase of the flight velocity will be more favorable than an increase of the lift coefficient. The arrangement of the control permits lateral stability to be achieved in every case; a minimum rolling moment due to sideslip proves of great help.
    Keywords: Aircraft Stability and Control
    Type: NACA-TM-1248 , ZWB Forschungsbericht; Rept-1819
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  • 33
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    Transonic Stability and Control Investigation of a 1/80-Scale Model of the Consolidated Vultee Skate 9 Seaplane, TED No. NACA DE 345: Transonic-Bump Method (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation of the longitudinal stability and of the all-movable horizontal tail and aileron control of a 1/80-scale reflection-plane model of the Consolidated Vultee Skate 9 seaplane has been made through a Mach number range of 0.6 to 1.16 on the transonic bump of the Langley high-speed 7- by 10-foot tunnel. At moderate lift coefficients (0.4 to 0.8) and below a Mach number of 1.0 the model was statically unstable longitudinally. The static longitudinal stability of the model at low lift coefficients increased with Mach number corresponding to a shift in aerodynamic center from 37 percent mean aerodynamic chord at a Mach number of 0.60 to 64 percent at a Mach number of 1.10. Estimates indicate that the tail deflection angle required for steady flight and for accelerated maneuvers of the Skate 9 airplane would probably not vary greatly with Mach number at sea level, but for accelerated maneuvers at altitude the tail deflection angle would probably vary erratically with Mach number. The variation of rolling-moment coefficient with aileron deflection angle was approximately linear, agreed well with theory, and held for the range of aileron deflections tested (-17.1 deg to 16.6 deg). At low lift coefficients the drag rise occurred at Mach numbers of 0.95 and 0.90 for the wing alone and the complete model, respectively. The effects of the canopy on the model were small. For the ranges investigated, angle-of-attack and Mach number changes caused no large pressure drops in the jet-engine duct.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL51E22
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  • 34
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    Yawed-Landing Investigation of a Model of the Convair Y2-2 Airplane, TED No. NACA DE 363 (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A model of the Convair Y2-2 airplane was tested in Langley tank no. 2 to determine whether satisfactory stability in yawed landings was possible with a certain ventral fin. Free-body landings were made in smooth and rough water at two speeds and two rates of descent with the model yawed 15deg. The behavior of the model was determined by visual observations and from motion-picture re.cords. It was concluded that satisfactory stability was possible with the ventral fin as tested but that the characteristics of the model shock absorbers and the settings of the elevon control surfaces had an appreciable influence on behavior.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL51H17A
    Format: application/pdf
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  • 35
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    Results of the Flight Test of a Dummy of the MX-656 Rocket-Propelled Models (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The data obtained from the flight of a simplified (dummy) rocket-propelled model of the MX-656 have been analyzed to determine the booster-model characteristics and the model-alone characteristics up to a Mach number of 1.3. The data indicate that the model-booster combination is satisfactory. The model alone is longitudinally stable i n the Mach number range covered by the test (0.9 to 1.3) with the center of gravity at -15 percent of the mean aerodynamic chord. With the stabilizer setting at 0 deg. the variation of normal-force coefficient with Mach number is not large. The total-drag-coefficient variation with Mach number is not unusual. About 12 percent of the total drag at a Mach number of 1.3 can be attributed to body base drag.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-SL50A07
    Format: application/pdf
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