Langley Aeronautical Laboratory
Overview
Works:  1,508 works in 2,082 publications in 1 language and 4,608 library holdings 

Classifications:  TL672, 629.1 
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.
Most widely held works about
Langley Aeronautical Laboratory
 Engineer in charge a history of the Langley Aeronautical Laboratory, 19171958 by James R Hansen( Book )
 Historical perspectives on thermostructural research at the NACA Langley Aeronautical Laboratory from 1948 to 1958 by Richard R Heldenfels( Book )
 Subcommittee hearing on H.R. 6336, to promote the national defense by authorizing the construction of aeronautical research facilities by the National Advisory Committee for Aeronautics necessary to the effective prosecution of aeronautical research by United States( Book )
 Full committee hearing on H.R. 6319, H.R. 5990, H.R. 6203, H.R. 6336, H.R. 4511 by United States( Book )
 From biplanes to Apollo : the NASA Langley Historic District by Joseph R Chambers( Book )
 by Robert R Gilruth( )
 Status of NACA research applicable to personal aircraft by United States( Book )
 by Melvin N Gough( )
 by Christopher C Kraft( )
 by Blake W Corson( )
 Papers by Christopher C Kraft( )
 NACALMAL and you by Langley Aeronautical Laboratory( Book )
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Most widely held works by
Langley Aeronautical Laboratory
Recent research on the determination of natural modes and frequencies of aircraft wing structures by
Langley Aeronautical Laboratory(
Book
)
2 editions published in 1956 in English and held by 31 WorldCat member libraries worldwide
2 editions published in 1956 in English and held by 31 WorldCat member libraries worldwide
Aerodynamic theory and its application to flutter by
Langley Aeronautical Laboratory(
Book
)
1 edition published in 1956 in English and held by 29 WorldCat member libraries worldwide
1 edition published in 1956 in English and held by 29 WorldCat member libraries worldwide
Note on the importance of imperfectgas effects and variation of heat capacities on the isentropic flow of gases by
Coleman duP Donaldson(
)
2 editions published in 1948 in English and held by 13 WorldCat member libraries worldwide
Abstract: The errors involved in using the perfectgas law pv=RT and the assumption of constant heat capacities are evaluated. The basic equations of gas flows taking into account these phenomena separately and at the same time are presented
2 editions published in 1948 in English and held by 13 WorldCat member libraries worldwide
Abstract: The errors involved in using the perfectgas law pv=RT and the assumption of constant heat capacities are evaluated. The basic equations of gas flows taking into account these phenomena separately and at the same time are presented
Comparison of yaw characteristics of a singleengine airplane model with singlerotating and dualrotating propellers by R. H Neely(
)
3 editions published in 1944 in English and held by 12 WorldCat member libraries worldwide
Tests were made n the NACA 29foot pressure tunnel to determine the yaw characteristics of a 0.32scale model of a singleengine, fightertype airplane with sixblade singlerotating and dualrotating propellers. The propellers used in the investigation were of the same solidity and plan form. Force and moment characteristics of the model, with the exception of the rollingmoment characteristics, are presented for several model and power conditions. Curves are given that show estimated ruddercontrol characteristics of the design airplane in steady sideslips
3 editions published in 1944 in English and held by 12 WorldCat member libraries worldwide
Tests were made n the NACA 29foot pressure tunnel to determine the yaw characteristics of a 0.32scale model of a singleengine, fightertype airplane with sixblade singlerotating and dualrotating propellers. The propellers used in the investigation were of the same solidity and plan form. Force and moment characteristics of the model, with the exception of the rollingmoment characteristics, are presented for several model and power conditions. Curves are given that show estimated ruddercontrol characteristics of the design airplane in steady sideslips
Heat transfer and skin friction for turbulent boundary layers on heated or cooled surfaces at high speeds by
Coleman duP Donaldson(
)
2 editions published in 1952 in English and held by 12 WorldCat member libraries worldwide
Abstract: The method presented in NACA TN 2692 for evaluating the skin friction of a turbulent boundary layer in compressible flow on an insulated surface is extended to evaluate the turbulent skin friction and heat transfer in compressible flow on a surface which is heated or cooled. The results of this analysis are in good agreement with the heat transfers measured in flight on the NACA RM10 missile up to Mach number of 3.8
2 editions published in 1952 in English and held by 12 WorldCat member libraries worldwide
Abstract: The method presented in NACA TN 2692 for evaluating the skin friction of a turbulent boundary layer in compressible flow on an insulated surface is extended to evaluate the turbulent skin friction and heat transfer in compressible flow on a surface which is heated or cooled. The results of this analysis are in good agreement with the heat transfers measured in flight on the NACA RM10 missile up to Mach number of 3.8
An electromagneticanalogy method of solving liftingsurfacetheory problems by
Robert S Swanson(
)
3 editions published in 1945 in English and held by 12 WorldCat member libraries worldwide
A method is suggested for making liftingsurface calculations by means of magnetic measurements of an electromagneticanalogy model. The method is based on the perfect analogy between the strength of the magnetic field around a conductor and the strength of the inducedvelocity field around a vortex. Electric conductors are arranged to represent the vortex sheet. The magneticfield strength is determined by measuring, with an electronic voltmeter, the voltage induced in a small search coil by the alternating current in the wires representing the vortex sheet. Solutions of nonlinear liftingsurface problems may be obtained by placing the conductors representing the trailing vortices along the fluid lines (Helmholtz condition). A potentialflow solution for the distortion and rolling up of the trailingvortex sheet may be obtained. By use of the PrandtlGlauert rule, the liftingsurface theory may be adapted to include firstorder compressibility effects
3 editions published in 1945 in English and held by 12 WorldCat member libraries worldwide
A method is suggested for making liftingsurface calculations by means of magnetic measurements of an electromagneticanalogy model. The method is based on the perfect analogy between the strength of the magnetic field around a conductor and the strength of the inducedvelocity field around a vortex. Electric conductors are arranged to represent the vortex sheet. The magneticfield strength is determined by measuring, with an electronic voltmeter, the voltage induced in a small search coil by the alternating current in the wires representing the vortex sheet. Solutions of nonlinear liftingsurface problems may be obtained by placing the conductors representing the trailing vortices along the fluid lines (Helmholtz condition). A potentialflow solution for the distortion and rolling up of the trailingvortex sheet may be obtained. By use of the PrandtlGlauert rule, the liftingsurface theory may be adapted to include firstorder compressibility effects
Analysis of factors affecting net lift increment attainable with trailingedge split flaps on tailless airplanes by
Marvin Pitkin(
)
in English and held by 12 WorldCat member libraries worldwide
An analysis has been made of factors affecting the net lift increment attainable with trailingedge split flaps on tailless airplanes. The flaps investigated in the analysis were designed to contribute zero pitching moments about the wing aerodynamic center when deflected. Calculations were made of the lift and pitchingmoment characteristics of flaps of this type over a range of design conditions in which sweepback angle, aspect ratio, taper ratio, flap chord, and flap deflection were widely varied. In addition, calculations were made to determine the effect of the various parameters upon the loss in lift incurred in trimming the stability moments of a tailless airplane. A method is given for roughly estimating the maximum lift coefficient of tailless airplanes
in English and held by 12 WorldCat member libraries worldwide
An analysis has been made of factors affecting the net lift increment attainable with trailingedge split flaps on tailless airplanes. The flaps investigated in the analysis were designed to contribute zero pitching moments about the wing aerodynamic center when deflected. Calculations were made of the lift and pitchingmoment characteristics of flaps of this type over a range of design conditions in which sweepback angle, aspect ratio, taper ratio, flap chord, and flap deflection were widely varied. In addition, calculations were made to determine the effect of the various parameters upon the loss in lift incurred in trimming the stability moments of a tailless airplane. A method is given for roughly estimating the maximum lift coefficient of tailless airplanes
Windtunnel investigation of effects of a pusher propeller on lift, profile drag, pressure distribution, and boundarylayer
transition of a flapped wing by
Carl A Sandahl(
)
4 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
Some of the effects of pusherpropeller operation on the aerodynamic characteristics of a flapped wing were measured in the Langley propellerresearch tunnel. The effects of propeller operation on the lift and profile drag of the wing, on pressure distribution, and on the position of boundarylayer transition were obtained. The results indicated that, at fixed angles of attack and with flaps deflected, the wing lift increased appreciably with increasing thrust coefficient. With flaps retracted, no appreciable increase in lift with increases in thrust coefficient was measured. Chordwise pressure distributions at several spanwise stations indicated that the effect of propeller operation was greatest in the region immediately ahead of the propeller and that the effect extended outboard from the propeller axis for about 2.5 propeller radii. Measurements of boundarylayer velocity on the forward part of the upper surface of the wing showed no appreciable shift of transition in the range of thrust coefficients investigated
4 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
Some of the effects of pusherpropeller operation on the aerodynamic characteristics of a flapped wing were measured in the Langley propellerresearch tunnel. The effects of propeller operation on the lift and profile drag of the wing, on pressure distribution, and on the position of boundarylayer transition were obtained. The results indicated that, at fixed angles of attack and with flaps deflected, the wing lift increased appreciably with increasing thrust coefficient. With flaps retracted, no appreciable increase in lift with increases in thrust coefficient was measured. Chordwise pressure distributions at several spanwise stations indicated that the effect of propeller operation was greatest in the region immediately ahead of the propeller and that the effect extended outboard from the propeller axis for about 2.5 propeller radii. Measurements of boundarylayer velocity on the forward part of the upper surface of the wing showed no appreciable shift of transition in the range of thrust coefficients investigated
Formulas for propellers in yaw and charts on the sideforce derivative by
H. S Ribner(
)
3 editions published in 1943 in English and held by 11 WorldCat member libraries worldwide
General formulas are given for propellers for the rate of change of sideforce coefficient with angle of yaw and for the rate of change of pitchingmoment coefficient with angle of yaw. Charts of the sideforce derivative are given for two propellers of different plan form. The charts cover solidities of two to six blades and single and dual rotation. The blade angles range from 15° or 20° to 60°. The equations, and the charts computed from the equations, are based on an unpublished analysis, which incorporates factors not adequately covered in previously published work and gives good agreement with experiment over a wide range of operating conditions. A study of the equations indicates that they are consistent with the following physical interpretation: In developing side force, the propeller acts like a fin of which the area is the projected side area of the propeller, the effective aspect ratio is of the order of 8, and the effective dynamic pressure is roughly that at the propeller disk as augmented by the inflow. The variation of the inflow velocity, for a fixedpitch propeller, accounts for most of the variation of side force with advancediameter ratio. The charts may be applied to obtain the rate of change of normalforce coefficient with angle of attack of the axis of rotation if proper account is taken of the upwash or downwash from the wing
3 editions published in 1943 in English and held by 11 WorldCat member libraries worldwide
General formulas are given for propellers for the rate of change of sideforce coefficient with angle of yaw and for the rate of change of pitchingmoment coefficient with angle of yaw. Charts of the sideforce derivative are given for two propellers of different plan form. The charts cover solidities of two to six blades and single and dual rotation. The blade angles range from 15° or 20° to 60°. The equations, and the charts computed from the equations, are based on an unpublished analysis, which incorporates factors not adequately covered in previously published work and gives good agreement with experiment over a wide range of operating conditions. A study of the equations indicates that they are consistent with the following physical interpretation: In developing side force, the propeller acts like a fin of which the area is the projected side area of the propeller, the effective aspect ratio is of the order of 8, and the effective dynamic pressure is roughly that at the propeller disk as augmented by the inflow. The variation of the inflow velocity, for a fixedpitch propeller, accounts for most of the variation of side force with advancediameter ratio. The charts may be applied to obtain the rate of change of normalforce coefficient with angle of attack of the axis of rotation if proper account is taken of the upwash or downwash from the wing
Numerical evaluation of the [epsilon]integral occurring in the Theodorsen arbitraryairfoil potential theory by
Irven Naiman(
)
3 editions published in 1944 in English and held by 11 WorldCat member libraries worldwide
A more precise method of evaluating the [epsilon]integral occurring in the arbitrary airfoil theory of Theodorsen (NACA Reps. nos. 411 and 452) is developed by retaining higher order terms in the Taylor expansion and by use of Simpson's rule. Formulas are given for routing calculation of the [epsilon]integral and for the necessary computational coefficients. The computational coefficients are tabulated for a 40point division of the range of integration from 0 to 2[pi]. With no increase in computational work the systematic error in the numerical value of [epsilon] is reduced from the order of 1 percent to approximately 0.1 percent
3 editions published in 1944 in English and held by 11 WorldCat member libraries worldwide
A more precise method of evaluating the [epsilon]integral occurring in the arbitrary airfoil theory of Theodorsen (NACA Reps. nos. 411 and 452) is developed by retaining higher order terms in the Taylor expansion and by use of Simpson's rule. Formulas are given for routing calculation of the [epsilon]integral and for the necessary computational coefficients. The computational coefficients are tabulated for a 40point division of the range of integration from 0 to 2[pi]. With no increase in computational work the systematic error in the numerical value of [epsilon] is reduced from the order of 1 percent to approximately 0.1 percent
Preliminary flight research on an allmovable horizontal tail as a longitudinal control for flight at high mach numbers by
Harold F Kleckner(
)
3 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
The NACA is conducting flight tests of an allmovable horizontal tail installed on a Curtiss XP42 airplane because of its possible advantages as a longitudinal control for flight at high Mach numbers. The results are presented for some preliminary tests in the lowspeed range for which the tail was very closely balanced aerodynamically and a bobweight was used to obtain stable stickforce variations with speed and acceleration. For these tests, the tail was hinged at 0.24 chord and was tried with two arrangements of servotab control. The elevator control was found to be unsatisfactory with the control arrangements tested. Although there were sufficient variation of stick force with acceleration in steady turns and a stable stickforce variation with speed, the nearzero variation of stick force with stick deflection resulted in an extremely sensitive control that required continuous attention in order to avoid motions of the airplane due to inadvertent movements of the control stick. For subsequent tests, the servotabs are being connected as geared unbalancing tabs in order that more conventional elevator hingemoment characteristics may be obtained. The expected advantages of the allmovable tail with a control system utilizing tabs would of course be limited to flight at Mach numbers below those for which severe compressibility effects are encountered on the tail itself. For higher Mach numbers, the allmovable tail would require an irreversible powerboost control in order to handle the large hingemoment increases that are expected
3 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
The NACA is conducting flight tests of an allmovable horizontal tail installed on a Curtiss XP42 airplane because of its possible advantages as a longitudinal control for flight at high Mach numbers. The results are presented for some preliminary tests in the lowspeed range for which the tail was very closely balanced aerodynamically and a bobweight was used to obtain stable stickforce variations with speed and acceleration. For these tests, the tail was hinged at 0.24 chord and was tried with two arrangements of servotab control. The elevator control was found to be unsatisfactory with the control arrangements tested. Although there were sufficient variation of stick force with acceleration in steady turns and a stable stickforce variation with speed, the nearzero variation of stick force with stick deflection resulted in an extremely sensitive control that required continuous attention in order to avoid motions of the airplane due to inadvertent movements of the control stick. For subsequent tests, the servotabs are being connected as geared unbalancing tabs in order that more conventional elevator hingemoment characteristics may be obtained. The expected advantages of the allmovable tail with a control system utilizing tabs would of course be limited to flight at Mach numbers below those for which severe compressibility effects are encountered on the tail itself. For higher Mach numbers, the allmovable tail would require an irreversible powerboost control in order to handle the large hingemoment increases that are expected
Windtunnel tests of dualrotating propellers with systematic differences in number of blades, blade setting, and rotational
speed of front and rear propellers by
W. H Gray(
)
3 editions published in 1944 in English and held by 11 WorldCat member libraries worldwide
The advent of dualrotating propellers has created a need for information concerning the effect of the number of blades of the front and rear propellers, relative rotational speeds, and small changes in the blade angles of the rear propeller. Results of aerodynamic tests of sevenblade propellers, which were considered as a possible arrangement to avoid vibration difficulties, are presented herein. Variations of relative blade angle and rotational speeds of the front and rear components of a sixblade dualrotating propeller were also investigated. The test program was an extension of previous work on dualrotating propellers at the NACA propellerresearch tunnel; the propeller blades and test body were those used in the previous tests. The results indicated that envelope efficiencies of a sevenblade propeller with three blades in the front hub and four in the rear were form 0 to 1 1/2 percent lower than envelope efficiencies for the sixblade dualrotating propeller; four blades in the front hub and three in the rear resulted in efficiencies 1/2 to 3 1/2 percent lower than those obtained with the sixblade propeller. This conclusion applies to bladeangle settings of the front and rear propellers to absorb equal power at peak efficiency when the rotational speeds were held equal
3 editions published in 1944 in English and held by 11 WorldCat member libraries worldwide
The advent of dualrotating propellers has created a need for information concerning the effect of the number of blades of the front and rear propellers, relative rotational speeds, and small changes in the blade angles of the rear propeller. Results of aerodynamic tests of sevenblade propellers, which were considered as a possible arrangement to avoid vibration difficulties, are presented herein. Variations of relative blade angle and rotational speeds of the front and rear components of a sixblade dualrotating propeller were also investigated. The test program was an extension of previous work on dualrotating propellers at the NACA propellerresearch tunnel; the propeller blades and test body were those used in the previous tests. The results indicated that envelope efficiencies of a sevenblade propeller with three blades in the front hub and four in the rear were form 0 to 1 1/2 percent lower than envelope efficiencies for the sixblade dualrotating propeller; four blades in the front hub and three in the rear resulted in efficiencies 1/2 to 3 1/2 percent lower than those obtained with the sixblade propeller. This conclusion applies to bladeangle settings of the front and rear propellers to absorb equal power at peak efficiency when the rotational speeds were held equal
On slenderbody theory at transonic speeds by
Keith C Harder(
)
2 editions published in 1954 in English and held by 11 WorldCat member libraries worldwide
Abstract: The basic ideas of the slenderbody approximation have been applied to the nonlinear transonicflow equation for the velocity potential in order to obtain some of the essential features of slenderbody theory at transonic speeds. The results of the investigation are presented from a unified point of view which demonstrates the similarity of slenderbody solutions in the various Mach number ranges. The transonic area rule and some conditions concerning its validity follow from the analysis
2 editions published in 1954 in English and held by 11 WorldCat member libraries worldwide
Abstract: The basic ideas of the slenderbody approximation have been applied to the nonlinear transonicflow equation for the velocity potential in order to obtain some of the essential features of slenderbody theory at transonic speeds. The results of the investigation are presented from a unified point of view which demonstrates the similarity of slenderbody solutions in the various Mach number ranges. The transonic area rule and some conditions concerning its validity follow from the analysis
Effects of compressibility on the maximum lift characteristics and spanwise load distribution of a 12footspan fightertype
wing of NACA 230series airfoil sections by E. O Pearson(
)
3 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
Force and pressuredistribution measurements were made on a fightertype wing model of conventional NACA 230series airfoil sections in the Langley 16foot highspeed tunnel to determine the effects of compressibility on the maximum lift characteristics and the spanwise load distribution. The range of angle of attack investigated was from 10° to 24°. The Mach number range was from 0.20 to 0.70 at small and medium angles of attack and from 0.15 to 0.625 at very large angles of attack. In the Mach number range from 0.15 to 0.55, the maximum lift coefficient first increased with increasing Mach number and then decreased rapidly after having reached a peak value at a Mach number of 0.30. At Mach numbers higher than 0.55, the rate of decrease of maximum lift coefficient with Mach number was considerably reduced. At these higher speeds the lift coefficient continued to increase with angle of attack well beyond the angle at which marked flow separation or stalling occurred, and the maximum lift coefficient was reached at angles 10° and 12° beyond the stalling angle. No significant changes in the span load distribution were found to occur below the stall at any of the test speeds. When the wing stalled at high speeds, the resultant load underwent a moderate outboard shift, which resulted in increases in root bending moment up to about 10 percent
3 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
Force and pressuredistribution measurements were made on a fightertype wing model of conventional NACA 230series airfoil sections in the Langley 16foot highspeed tunnel to determine the effects of compressibility on the maximum lift characteristics and the spanwise load distribution. The range of angle of attack investigated was from 10° to 24°. The Mach number range was from 0.20 to 0.70 at small and medium angles of attack and from 0.15 to 0.625 at very large angles of attack. In the Mach number range from 0.15 to 0.55, the maximum lift coefficient first increased with increasing Mach number and then decreased rapidly after having reached a peak value at a Mach number of 0.30. At Mach numbers higher than 0.55, the rate of decrease of maximum lift coefficient with Mach number was considerably reduced. At these higher speeds the lift coefficient continued to increase with angle of attack well beyond the angle at which marked flow separation or stalling occurred, and the maximum lift coefficient was reached at angles 10° and 12° beyond the stalling angle. No significant changes in the span load distribution were found to occur below the stall at any of the test speeds. When the wing stalled at high speeds, the resultant load underwent a moderate outboard shift, which resulted in increases in root bending moment up to about 10 percent
The flow past a straight and a sweptwingbody combination and their equivalent bodies of revolution at Mach numbers near
1.0 by
W. F Lindsey(
)
2 editions published in 1954 in English and held by 11 WorldCat member libraries worldwide
Abstract: The complete flow fields past a straight and a swept wingbody combination and their equivalent bodies of revolution at Mach numbers around 1.0 have been observed by means of motionpicture schlieren photography. The results of these observations indicate that the shock growth and positions on the wingbody combinations are closely reproduced in the flow past their respective equivalent bodies
2 editions published in 1954 in English and held by 11 WorldCat member libraries worldwide
Abstract: The complete flow fields past a straight and a swept wingbody combination and their equivalent bodies of revolution at Mach numbers around 1.0 have been observed by means of motionpicture schlieren photography. The results of these observations indicate that the shock growth and positions on the wingbody combinations are closely reproduced in the flow past their respective equivalent bodies
Liftingsurfacetheory values of the damping in roll and of the parameter used in estimating aileron stick forces by
Robert S Swanson(
)
3 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
An investigation was made by liftingsurface theory of a thin elliptic wing of aspect ratio 6 in a steady roll by means of the electromagneticanalogy method. From the results, aspectratio corrections for the damping in roll and aileron hinge moments for a wing in steady roll were obtained that are considerably more accurate than those given by liftingline theory. Firstorder effects of compressibility were included in the computations. The results obtained by liftingsurface theory indicate that the damping in roll for a wing of aspect ratio 6 is 13 percent less than that given by liftingline theory and 5 percent less than that given by liftingline theory with the edgevelocity correction derived by Robert T. Jones applied. The results are extended to wings of other aspect ratios
3 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
An investigation was made by liftingsurface theory of a thin elliptic wing of aspect ratio 6 in a steady roll by means of the electromagneticanalogy method. From the results, aspectratio corrections for the damping in roll and aileron hinge moments for a wing in steady roll were obtained that are considerably more accurate than those given by liftingline theory. Firstorder effects of compressibility were included in the computations. The results obtained by liftingsurface theory indicate that the damping in roll for a wing of aspect ratio 6 is 13 percent less than that given by liftingline theory and 5 percent less than that given by liftingline theory with the edgevelocity correction derived by Robert T. Jones applied. The results are extended to wings of other aspect ratios
Some notes on the determination of the stickfree neutral point from windtunnel data by
Marvin Schuldenfrei(
)
3 editions published in 1944 in English and held by 11 WorldCat member libraries worldwide
The effect on static longitudinal stability of freeing the elevator is shown to be similar to the effect of altering the slope of the tail life curve by a factor that depends upon the aerodynamic characteristics of the horizontal tail surfaces. The stickfree neutral point may then be determined from stickfixed data by taking account of the reduction of tail effectiveness. Two graphical methods for determining the stickfree neutral point, which are extensions of the methods commonly used to determine the stickfixed neutral point, are presented. A mathematical formula for computing the stickfree neutral point is also given. These methods may be applied to determine approximately the increase in tail size necessary to shift the neutral point (with stick free or fixed) to any desired location on an airplane having inadequate longitudinal stability
3 editions published in 1944 in English and held by 11 WorldCat member libraries worldwide
The effect on static longitudinal stability of freeing the elevator is shown to be similar to the effect of altering the slope of the tail life curve by a factor that depends upon the aerodynamic characteristics of the horizontal tail surfaces. The stickfree neutral point may then be determined from stickfixed data by taking account of the reduction of tail effectiveness. Two graphical methods for determining the stickfree neutral point, which are extensions of the methods commonly used to determine the stickfixed neutral point, are presented. A mathematical formula for computing the stickfree neutral point is also given. These methods may be applied to determine approximately the increase in tail size necessary to shift the neutral point (with stick free or fixed) to any desired location on an airplane having inadequate longitudinal stability
Propellers in yaw by
H. S Ribner(
)
3 editions published in 1943 in English and held by 11 WorldCat member libraries worldwide
It was realized as early as 1909 that a propeller in yaw develops a side force like that of a fin. In 1917, R.G. Harris expressed this force in terms of the torque coefficient for the unyawed propeller. Of several attempts to express the side force directly in terms of the shape of the blades, however, none has been completely satisfactory. An analysis that incorporates induction effects not adequately covered in previous work and that gives good agreement with experiment over a wide range of operating conditions is presented herein. The present analysis shows that the fin analogy may be extended to the form of the sideforce expression and that the effective fin area may be taken as the projected side area of the propeller. The effective aspect ratio is of the order of 8 and the appropriate dynamic pressure is roughly that at the propeller disk as augmented by the inflow. The variation of the inflow velocity, for a fixedpitch propeller, accounts for most of the variation of side force with advancediameter ration V/nD
3 editions published in 1943 in English and held by 11 WorldCat member libraries worldwide
It was realized as early as 1909 that a propeller in yaw develops a side force like that of a fin. In 1917, R.G. Harris expressed this force in terms of the torque coefficient for the unyawed propeller. Of several attempts to express the side force directly in terms of the shape of the blades, however, none has been completely satisfactory. An analysis that incorporates induction effects not adequately covered in previous work and that gives good agreement with experiment over a wide range of operating conditions is presented herein. The present analysis shows that the fin analogy may be extended to the form of the sideforce expression and that the effective fin area may be taken as the projected side area of the propeller. The effective aspect ratio is of the order of 8 and the appropriate dynamic pressure is roughly that at the propeller disk as augmented by the inflow. The variation of the inflow velocity, for a fixedpitch propeller, accounts for most of the variation of side force with advancediameter ration V/nD
Use of variableratio geared tabs to improve stickforce characteristics in turning flight by
Harold F Kleckner(
)
3 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
In flight tests of an experimental elevator with geared tabs, a cockpit control over the tab gear ratio was found to be satisfactory for adjusting the stick force per g in turning flight according to the pilot's preference. This type of control appears to have application for increasing the centerofgravity range for satisfactory stick forces in turning flight. Sample calculations made for a fighter airplane indicated that satisfactory stick forces in turning flight can be obtained for any centerofgravity position at which the elevator control meets other requirements
3 editions published in 1945 in English and held by 11 WorldCat member libraries worldwide
In flight tests of an experimental elevator with geared tabs, a cockpit control over the tab gear ratio was found to be satisfactory for adjusting the stick force per g in turning flight according to the pilot's preference. This type of control appears to have application for increasing the centerofgravity range for satisfactory stick forces in turning flight. Sample calculations made for a fighter airplane indicated that satisfactory stick forces in turning flight can be obtained for any centerofgravity position at which the elevator control meets other requirements
A summary of available knowledge concerning skin friction and heat transfer and its application to the design of highspeed
missiles by
Morris W Rubesin(
)
2 editions published in 1951 in English and held by 11 WorldCat member libraries worldwide
Abstract: A review is made of the existing information concerning boundarylayer characteristics: the temperature recovery, the skinfriction coefficients, the heattransfer coefficients of both the laminar and turbulent boundary layers and the position of the transition from laminar to turbulent flow. Comparison is made between existing flight data and results computed by the boundarylayer momentumintegral method in a preliminary attempt to establish some rational way of approaching the design of a missile whose Mach number range and body geometry are markedly different from those of existing data
2 editions published in 1951 in English and held by 11 WorldCat member libraries worldwide
Abstract: A review is made of the existing information concerning boundarylayer characteristics: the temperature recovery, the skinfriction coefficients, the heattransfer coefficients of both the laminar and turbulent boundary layers and the position of the transition from laminar to turbulent flow. Comparison is made between existing flight data and results computed by the boundarylayer momentumintegral method in a preliminary attempt to establish some rational way of approaching the design of a missile whose Mach number range and body geometry are markedly different from those of existing data
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Related Identities
 United States National Advisory Committee for Aeronautics
 Hansen, James R. Author
 United States National Aeronautics and Space Administration Scientific and Technical Information Office
 Langley Research Center
 Heldenfels, Richard R. Author
 Theodorsen, Theodore Author
 United States National Aeronautics and Space Administration Scientific and Technical Information Branch
 Zalovcik, John A. Author
 United States National Advisory Committee for Aeronautics Advisory Group for Aeronautical Research and Development
 Donaldson, Coleman duP Author
Associated Subjects
Acquisition of property Aerodynamics, Transonic AerodynamicsResearch Aerofoils Aeronautics, MilitaryResearch AeronauticsResearch Ailerons Airplanes, Military Airplanes, Tailless AirplanesWings Armed Forces (United States) Armed ForcesAppointments and retirements Armed ForcesOfficials and employees Body of revolution CaliforniaSan Francisco CaliforniaSan FranciscoPresidio of San Francisco Charts, diagrams, etc Civil defense Compressibility Conveyancing Elevators (Airplanes) Engineers Fighter planes Flaps (Airplanes) Flutter (Aerodynamics) Gases, Real Hancock, Joy Bright, HawaiiKahului Isentropic expansion Land titlesRegistration and transfer Langley Aeronautical Laboratory Lift (Aerodynamics) Mach number Military research New York (State)Schenectady OhioCleveland Propellers, Aerial Propellers, AerialTesting Propulsion Laboratory (U.S.) Research Skin friction (Aerodynamics) Thermal stressesResearch United States United States.National Advisory Committee for Aeronautics United States.Naval Reserve.Women's Reserve United States.Navy United States.Navy Department Vibration (Aeronautics) VirginiaHampton Yawing (Aerodynamics)
Alternative Names
Hampton (Va.). Langley Aeronautical Laboratory
Langley Air Force Base (Va.). Langley Aeronautical Laboratory
Langley Memorial Aeronautical Laboratory
United States. Langley Aeronautical Laboratory
United States. Langley Aeronautical Laboratory, Hampton, Va.
United States. Langley Memorial Aeronautical Laboratory
United States. National Advisory Committee for Aeronautics. Langley Aeronautical Laboratory
USA National Advisory Committee for Aeronautics Langley Memorial Aeronautical Laboratory
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