Research abstracts

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Title:
Research abstracts
Physical Description:
93 v. : ; 27 cm.
Language:
English
Creator:
United States -- National Advisory Committee for Aeronautics
Publisher:
National Advisory Committee for Aeronautics
Place of Publication:
Washington, D.C
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irregular
completely irregular

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Subjects / Keywords:
Aeronautics -- Abstracts -- Periodicals   ( lcsh )
Aeronautics -- Research -- Abstracts -- Periodicals   ( lcsh )
Genre:
serial   ( sobekcm )
federal government publication   ( marcgt )
abstract or summary   ( marcgt )

Notes

Statement of Responsibility:
National Advisory Committee for Aeronautics.
Dates or Sequential Designation:
Abstracts no. 1 (June 15, 1951)-no. 93 (Nov. 30, 1955).

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 001469326
notis - AGY1019
oclc - 01471285
lccn - 86657025
issn - 0499-9274
Classification:
lcc - TL501 .U5895
System ID:
AA00009235:00038

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National Advisory Committee for Aeronautics


Research Abstracts


DECEMBER 21, 1954


CURRENT NACA REPORTS





i' A


LA Tm rf o


CONTRIBUTION TO THE THEORY OF TAIL-
WHEEL SHIMMY. (Beitrag zur Theorie des
Spornradflatterns). M. Melzer. December 1954.
37p. diagrs., 7 tabs. (NACA TM 1380. Trans. from
Focke-Wulf Flugzeugbau G.m.b.H., Bremen,
Technische Berlchte, v. 7, no. 2, 1940)

A basic theoretical and experimental investigation is
made of the shimmy behavior of a swiveling landing
gear, the experimental tests being conducted with a
small wheel mounted over a continuous belt. Effects
of wheel loading, rolling velocity, rearward position
of the wheel with respect to the swivel axis, tire
elasticity, and torsional flexibility of the fuselage are
investigated both experimentally and theoretically.
A major theoretical conclusion is that the motion of
a landing gear moving in a straight line without
fuselage elasticity is stable for a sufficiently large
rearward position of the wheel behind the swivel
axis, and this conclusion is well verified
quantitatively by the experimental data.



NACA TN 3240

THEORETICAL CALCULATIONS OF THE LATERAL
STABILITY DERIVATIVES FOR TRIANGULAR VER-
TICAL TAILS WITH SUBSONIC LEADING EDGES
TRAVELING AT SUPERSONIC SPEEDS. Percy J.
Bobbitt. December 1954. 68p. diagrs., photos.,
5 tabs. (NACA TN 3240)

Pressure-distribution expressions and stability deriv-
atives have been derived for zero-end-plate triangu-
lar vertical tails performing yawing, rolling, and
constant-lateral-acceleration motions by a method
for solving supersonic-conical flow boundary-value
problems. In addition, by using the yawing and
constant-lateral-acceleration results, the damping of
the vertical tail oscillating laterally in yaw has been
approximated to the first order in frequency. End-
plate effects have been discussed and suggestions
made to aid in their evaluation. In this connection
the complete-end-plate and no-end-plate stability
derivatives for the sideslip motion obtained from
other sources have been considered. The pressure-
distribution expressions and stability derivatives
presented are valid for a range of Mach numbers for
which the leading edge is subsonic and the trading
edge supersonic.


I


,--- "


*AVAILABLE ON LOAN ONLY. L -. .'
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 15i2 H ST, NW., WASHINGTON 2s, D C, CITING CODE NUMBER KBOVE EACH TITLE;
THE REPORT TITLE AND AUTHOR.
6AL9. J3092

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NO.75


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NACA TN 3251

A-THEORETICAL INVESTIGATION OF THE
..-,RORT-PERIOD DYNAMIC LONGITUDINAL
STABILITY OF AIRPLANE CONFIGURATIONS
HAVING ELASTIC WINGS OF 00 TO 600 SWEEP-
BACK. Milton D. McLaughhn. December 1954.
39p. diagrs., 2 Labs. (NACA TN 3251)

An analytical investigation utilizing the semirigid
approach was made to determine the effects of wing
flexibility on the dynamic longitudinal stability of
uhin-wing airplane configurations. The configura-
tions were assumed to vary in wing sweep angle
from 00 to 600, in center-of-gravity location Irom
25 percent mean aerodynamic chord to 45 percent
mean aerodynamic chord, and in ratio of wing mass
to airplane mass from 0.15 to 0.50. Solution were
obtained in the form of frequency and dampint of the
flexible wing mode and of the airplane short-period
stability mode. Results are compared with tse
obtained by simplified methods.


NACA TN 3267

BOUNDARY-LAYER TRANSITION AT MACH 3.12
WITH AND WITHOUT SINGLE ROUGHNESS
ELEMENTS. Paul F. Brnich. December 1954.
41p. diagrs. (NACA TN 3267)

Transition on a hollow cylinder alined with the air
stream was studied in a Mach 12 wind tunnel at
Reynolds numbers from I x 10 to 7 x 10 per inch.
Transition was observed from surface-temperature
distributions and schlieren photographs with and
without single roughness elements. Without
roughness, the transition Reynolds number varied as
the square root of the stream Reynolds number per
inch. The results obtained with roughness followed
the trends of Dryden's low-speed correlation. The
degree of roughness required was found to be three
to seven times as great as for low-speed flows.



NACA TN 3280

ELECTRICAL ANALOGIES FOR STIFFENED
SHELLS WITH FLEXIBLE RINGS. R. H. MacNeal,
California Institute of Technology. December 1954.
35p. diagrs., 2 tabs. (NACA TN 3280)

First, an electrical analogy is developed for a
circular shell with a straight axis and variable 0
radius. This analogy is extended to noncircular-
cylinders. Then an electrical analogy is derived for .
.. l.


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+!i.
h, ..





NACA
RESEARCH ABSTRACTS NO. 75


rings with variable radii of curvature; a simplified
circuit for circular rings is also presented. The
simplifications that occur when the rings are
assumed to be rigid are discussed. Finally, results
are given for two sample problems solved on an
analog computer. The second problem concerns a
cantilever conical shell and illustrates the manner
in which the shell and ring circuits are intercon-
nected.


NACA TN 3288

ON THE ANALYSIS OF LINEAR AND NONLINEAR
DYNAMICAL SYSTEMS FROM TRANSIENT-
RESPONSE DATA. Marvin Shinbrot. December
1954. 51p. diagrs., 6 tabs. (NACA TN 3288)

A general theory of the so-called "equations-of-
motion" methods for linear dynamical systems is
developed first. It is then shown that when viewed
from this general point of vantage, all of these
linear methods can be extended in a perfectly
natural way to apply to nonlinear systems. In
addition, through use of this theory, a new method,
possessing certain advantages over previous
methods, is derived. Application of this method is
demonstrated by several examples.



NACA TN 3301

COMPARISON OF FLUTTER CALCULATIONS USING
VARIOUS AERODYNAMIC COEFFICIENTS WITH
EXPERIMENTAL RESULTS FOR SOME RECTANGU-
LAR CANTILEVER WINGS AT MACH NUMBER 1.3.
Herbert C. Nelson and Ruby A. Rainey. November
1954. 22p. diagrs., 2 tabs. (NACA TN 3301)

A general Rayleigh analysis is used as a basis for
developing four methods of flutter analysis that are
applied to twelve rectangular low-aspect-ratio wings
with aspect ratios ranging from 3.00 to 4.55. These
wings were previously tested at a Mach number of
1.3 by progressively varying certain wing parameters
until flutter occurred. The four methods of flutter
analysis used are: section coefficients for harmoni-
cally pitching and translating rectangular wings in a
Rayleigh type of analysis, two-dimensional coeffi-
cients in a Rayleigh type of analysis, total coeffi-
cients for harmonically pitching and translating rec-
tangular wings in a representative-section analysis.
and two-dimensional coefficients in a representative-
section analysis. Each of the four methods involves
two degrees of freedom, namely, first bending and
first torsion of a cantilever wing. Comparison of the
analytical and experimental results indicates that the
use of aerodynamic coefficients derived on the basis
of three-dimensional flow leads to a significant
improvement in the correlation of theory and
experiment.



NACA TN 3304

INVESTIGATION OF THE AERODYNAMIC CHARAC-
TERISTICS OF A MODEL WING-PROPELLER COM-
BINATION AND OF THE WING AND PROPELLER
SEPARATELY AT ANGLES OF ATTACK UP TO 900.
John W. Draper and Richard E. Kuhn. November
1954. 72p. diagrs., photos., tab. (NACA TN 3304)


Results are presented of a wind-tunnel investigation
of the effects of slipstream for two large-diameter
propellers on the aerodynamic characteristics of a
wing model. The investigation covered angles of
attack from -100 to 900 and thrust coefficients repre-
senting free-stream velocities from zero to the nor-
mal range of cruising flight. An appreciable increase
in the angle of attack for maximum lift with increas-
ing power is indicated. A modification of the method
of Smelt and Davies for estimating the effects of
slipstreams on the lift-curve slope is presented.
Performance calculations for an assumed vertical
take-off airplane are also included. Forces and
moments on the wing-propeller combination and on
the wing and propellers separately are included. A
large nose-up pitching moment on the propeller
itself was found at high angles of attack.


NACA TN 3306

AN INVESTIGATION OF A LIFTING 10-PERCENT-
THICK SYMMETRICAL DOUBLE-WEDGE AIRFOIL
AT MACH NUMBERS UP TO 1. Milton D.
Humphreys. November 1954. 35p. diagrs., photos.,
tab. (NACA TN 3306)

Subsonic and transonic pressure-distribution and
drag-coefficient measurements on a two-dimensional
10-percent-thick symmetrical double-wedge airfoil
for a range of angles of attack are presented and
compared with theory and existing experimental data.


NACA TN 3307

AN INVESTIGATION OF A WING-PROPELLER CON-
FIGURATION EMPLOYING LARGE-CHORD PLAIN
FLAPS AND LARGE-DIAMETER PROPELLERS FOR
LOW-SPEED FLIGHT AND VERTICAL TAKE-OFF.
Richard E. Kuhn and John W. Draper. December
1954. 94p. diagrs., photos. (NACA TN 3307)

An investigation of the effectiveness of a wing
equipped with large-chord plain flaps and auxiliary
vanes in rotating the effective thrust vector from two
large-diameter propellers to a near vertical attitude
for vertical take-off and low-speed flight has been
conducted in the Langley 300 mph 7- by 10-foot tun-
nel. The model consisted of a semispan wing
equipped with a 60-percent-chord and a 30-percent-
chord plain flap. Two large-diameter overlapping
propellers, driven by electric motors, were used.
The tests covered a range of angle of attack from 00
to 900 and a thrust-coefficient range representative
of free-flight velocities from zero to the normal
range of cruising velocities.


NACA TN 3309

MECHANICAL PROPERTIES AT ROOM TEMPERA-
TURE OF FOUR CERMETS OF TUNGSTEN CAR-
BIDE WITH COBALT BINDER. Aldie E. Johnson,
Jr. December 1954. 16p. diagrs., photo., tab.
(NACA TN 3309)

Room-temperature stress-strain curves are pre-
sented for compression, tension, and shear loadings
on four compositions of tungsten carbide with cobalt
binder. Values of modulus of elasticity, modulus of
rigidity, Poisson's ratio in the elastic region, ulti-
mate strength, density, and hardness for the four
materials are tabulated.





NACA
RESEARCH ABSTRACTS NO. 75

NACA TN 3312

INITIAL EXPERIMENTS ON FLUTTER OF UN-
SWEPT CANTILEVER WINGS AT MACH NUMBER
1.3. W. J. Tuovila, John E. Baker and Arthur A.
Regier. November 1954. 19p. diagrs., photos.,
2 tabs. (NACA TN 3312. Formerly RM L8J11)

This paper contains experimental data for the bend-
ing torsion flutter of unswept cantilever wings at
Mach number of 1.3. Comparison with the theory of
flutter at supersonic speeds indicates that the theory
is applicable in predicting cantilever-wing flutter.
These preliminary experiments consider also various
overall effects of center-of-gravity location, elastic-
axis location, mass-density parameter, section
shape, and some wing-tip moments of inertia.




NACA TN 3316

SOME MEASUREMENTS OF NOISE FROM THREE
SOLID-FUEL ROCKET ENGINES. Leslie W.
Lassiter and Robert H. Heltkotter. December 1954.
21p. diagrs. (NACA TN 3316)

Results of a systematic experimental investigation of
the sound field of one type of rocket engine are given.
These comprise frequency spectra and overall pres-
sure data as obtained along a semicircle of 50-foot
radius about the engine. The spectra are shown to
be of random nature and the angular distribution is
shown to have a maximum at an angle of about 45o
off the jet axis downstream of the nozzle. Frequency
spectra and overall pressure magnitudes are also
given for a few isolated points in the sound fields of
two other solid-fuel rockets.




NACA TN 3320

FLIGHT MEASUREMENTS OF DRAG AND BASE
PRESSURE OF A FIN-STABILIZED PARABOLIC
BODY OF REVOLUTION (NACA RM-10) AT DIF-
FERENT REYNOLDS NUMBERS AND AT MACH
NUMBERS FROM 0.9 TO 3.3. H. Herbert Jackson,
Charles B. Rumsey and Leo T. Chauvin. November
1954. 20p. diagrs., photos. (NACA TN 3320.
Formerly RM L50G24)

Free-flght tests at supersonic speeds have been
made to determine the Reynolds number effects on
total drag and base drag of a fin-stabilized parabolic-
arc body of revolution having a body fineness ratio of
12.2 and designated the NACA RM-10 configuration.
The Reynolds number range of 14 x 106 to 210 x 106
was obtained by testing full-scale and half-scale
models through the Mach number range from 0.9
to 3.3.




NACA TN 3328

THEORETICAL AND EXPERIMENTAL INVESTIGA-
TION OF HEAT TRANSFER BY LAMINAR NATURAL
CONVECTION BETWEEN PARALLEL PLATES.
A. F. Lietzke. December 1954. 23p. diagrs.
(NACA TN 3328)


For the experimental work, parallel walls (one
heated uniformly and the other cooled uniformly)
were simulated by an annulus with an inner-to-
outer diameter ratio near 1. The theoretical
results are presented in equations for the velocity
and temperature profiles and the ratio of actual
temperature drop across the fluid to temperature
drop for pure conduction. No experimental
measurements were made of the velocity and
temperature profiles, but the experimental results
are compared with theory on the basis of the ratio
of the actual temperature drop to the temperature
drop for pure conduction. Good agreement was
obtained between theory and experiment for axial
temperature gradients above 400 F/ft.

NACA TN 3333

CORROSION OF METALS OF CONSTRUCTION BY
ALTERNATE EXPOSURE TO LIQUID AND
GASEOUS FLUORINE. Richard M. Gundzik and
Charles E. Feller. December 1954. 10p. photos.,
3 tabs. (NACA TN 3333)

The corrosion of 3S-0 and 52S-0 aluminum,
AISI 347 and 321 stainless steels, 'A' nickel, and
low-leaded brass by alternate exposure to liquid
and gaseous fluorine has been determined for
periods of up to 3-1/2 months. Under the conditions
of the experiments, the corrosion of these metals
was found to be negligible.


NACA TN 3341

AN ANALYTICAL ESTIMATION OF THE EFFECT
OF TRANSPIRATION COOLING ON THE HEAT-
TRANSFER AND SKIN-FRICTION CHARACTER-
ISTICS OF A COMPRESSIBLE TURBULENT
BOUNDARY LAYER. Morris W. Rubesin.
December 1954. 56p. diagrs. (NACA TN 3341)

An analysis based on mixing-length theory is
presented which indicates that surface blowing
associated with transpiration cooling systems
produces large reductions in both the heat-transfer
and skin-friction coefficients for a turbulent
boundary layer on a flat plate. The numerical
results are restricted to the case of air blowing
into air. The effects of blowing are indicated to be
similar for high-speed, compressible flow to those
for low-speed, incompressible flow.


NACA TN 3344

THEORETICAL AND EXPERIMENTAL INVESTIGA-
TION OF AERODYNAMIC-HEATING AND ISOTHER-
MAL HEAT-TRANSFER PARAMETERS ON A
HEMISPHERICAL NOSE WITH LAMINAR BOUNDARY
LAYER AT SUPERSONIC MACH NUMBERS.
Howard A. Stine and Kent Wanlass. December 1954.
48p. diagrs., photo., tab. (NACA TN 3344)

Measurements of the heat-transfer parameter,
Nu/R-e, based on local flow conditions just outside
the boundary layer and length of boundary-layer run
were carried out on the heated, isothermal surface
of a hemisphere-cylinder for a Mach number of 1.97
and Reynolds numbers eased on body diameter from
0.60 x 10 to 2.28 x 10 An approximate method
for calculating the distribution of heat-transfer
parameter, applicable to any body of revolution, was
developed which agrees well with the experimental
results.





NACA
RESEARCH ABSTRACTS NO. 75


NACA TN 3348

A SYSTEM FOR MEASURING THE DYNAMIC LAT-
ERAL STABILITY DERIVATIVES IN HIGH-SPEED
WIND TUNNELS. Henry C. Lessing, Thomas B.
Fryer and Merrill H. Mead. December 1954. 42p.
diagrs., photo. (NACA TN 3348)

A two-degree-of-freedom system in which rolling
oscillations are forced is described. Details of the
system, the theory, and the method of operation are
discussed. Tests are described in which certain of
the aerodynamic derivatives were simulated, and
the results of these tests, as well as the results of a
wind-tunnel investigation of a simple model, are
presented.



BRITISH REPORTS


N-34221 *

Forest Products Research Lab. (Gt.Brit.)
INVESTIGATIONS INTO GLUES AND GLUING.
PROGRESS REPORT EIGHTY APRIL 1954. THE
DURABILITY OF ASSEMBLY GLUES SERIES A.
REPORT ON THE PLY-CLAD FRAMES AFTER
SEVEN AND A HALF YEAR'S EXPOSURE. R. A. G.
Knight, L. S. Doman and G. E. Soane. 12p.,
3 tabs. (Supersedes Progress Reports 47 and 64)
(Forest Products Research Lab.)

An investigation is being conducted of the long-term
durability of all types of plywood glues tested under
conditions ranging from normal indoor storage to
full exposure. Results for a 7-1/2 year period are
presented.


N-34234 *

Aeronautical Research Council (Gt. Brit.)
THE APPLICATION OF CAMBER AND TWIST TO
SWEPT WINGS IN INCOMPRESSIBLE FLOW. G. G.
Brebner. 1954. 59p. diagrs., 2 tabs. (ARC CP 171)

This report provides a method of calculating the
effect of a certain type of camber on the chordwise
and spanwise loading of any wing in incompressible
potential flow. A detailed description is given of the
method of calculating the loading on a given cambered
swept wing. The design of wings incorporating
camber and twist to produce required chordwise and
spanwise loadings is treated in detail. It is shown
that at any spanwise position only one particular
combination of camber and incidence will give a
required chordwise loading.


N-34235*

Aeronautical Research Council (Gt. Brit.)
A NOTE ON THE BOUNDARY LAYER AND
STALLING CHARACTERISTICS OF AEROFOILS.
D. D. Carrow. 1954. 20p. diagrs. (ARC CP 174)
A qualitative explanation is suggested for the changes
which occur in the stalling characteristics of airfoils
as the Reynolds number is increased. This expla-
nation is based on the variation, with Reynolds


number, in the state of the boundary layer along the
upper surface at just below the stalling incidence.
Examples are given of the effect of Reynolds number
and of leading edge roughness on the stalling char-
acteristics of several airfoils, the changes from one
type of stall to another being explained, in each
case, by an alteration in the type of transition
region. Five major types of stall are explained
according to the manner in which the flow separation
at the stall develops.


N-34322*

Aeronautical Research Council (Gt. Brit.)
STRESS CONCENTRATIONS AT A CUT-OUT IN A
SWEPT WING. E. H. Mansfield. 1954. 21p.
diagrs., 5 tabs. (ARC R & M 2823; ARC 14,249.
Formerly RAE Structures 114)

The stress concentrations are determined for a
panel, bounded by main load-carrying members and
an oblique edge, such as might occur at a cut-out in
a swept wing. The solutions given are exact and
cover the effects of a member along the oblique edge
and of closely spaced stringers attached to the panel.


N-34323*

Aeronautical Research Council (Gt. Brit.)
FLIGHT TRIALS OF A ROCKET-PROPELLED
TRANSONIC RESEARCH MODEL: THE R. A. E. -
VICKERS ROCKET MODEL, PARTS I TO IV. Staff
of the Supersonics Division, Flight Section. 1954.
63p. diagrs., photos., 2 tabs. (ARC R & M 2835;
ARC 13,658. Formerly RAE Aero 2357)

The main body of the report contains a historical
review of the project, brief descriptions of the test
vehicles and experimental techniques,and the
detailed analysis of the final successful flight trial.
It is concluded that the method is an extremely dif-
ficult one and that the results it gives do not justify
the effort it demands. Detailed descriptions of the
test vehicle and its ancillary equipment are given in
appendices; the work leading up to the final
trial and the experimental methods and techniques
are described. The model was air-launched at
36,000 feet above a ground radar station.


N-34324*

Aeronautical Research Council (Gt. Brit.)
THE FREQUENCY RESPONSE OF THE ORDINARY
ROTOR BLADE, THE HILLER SERVO-BLADE,
AND THE YOUNG-BELL STABILIZER. G. J.
Sissingh. 1954. 19p. diagrs., 2 tabs. (ARC
R & M 2860; ARC 13,592. Formerly RAE Aero2367)

The report deals with the frequency response of
pivoted gyratory systems, and in particular, with
the response of the ordinary rotor blade, the Hiller
servo-blade, and the Young-Bell stabilizer to
sinusoidal disturbances caused by pitching oscilla-
tions with constant amplitude. Explicit formulas are
given for each individual system and the effect of the
frequency ratio on the control characteristics of the
Bell stabilizer and the Hiller servo-blade is shown
by vector loci.





NACA
RESEARCH ABSTRACTS NO. 75

N-34325'

Aeronautical Research Council (Gt. Brit.)
WATER AND AIR PERFORMANCE OF SEAPLANE
HULLS AS AFFECTED BY FAIRING AND FINENESS
RATIO. A. G. Smith and J. E. Allen. 1954. 27p.
diagrs., 5 tabs. (ARC R & M 2896; ARC 13,644.
Formerly MAEE F/Res/219)

An analysis is made of data on the variation of hull
air drag with length-beam ratio and degree of local
fairing, and of the maximum beam loading permitted
for reasonable hydrodynamic performance. The
effect of length-beam ratio on hull structure weight
is also discussed. Charts are given to help in the
selection of the best hull form for any given design
conditions. From these may be estimated dimen-
sions, air drag, and the maximum beam loading to
give a reasonable standard of water performance.

N-34326*

Aeronautical Research Council (Gt. Brit.)
A NOTE ON THE DYNAMIC STABILITY OF AIR-
CRAFT AT HIGH-SUBSONIC SPEEDS WHEN CON-
SIDERING UNSTEADY FLOW. W. J. G. Pinsker.
1954. 28p. diagrs., 7 tabs. (ARC R & M 2904;
ARC 13,567. Formerly RAE Aero 2378)

The effect of an increase in speed relative to the
speed of sound on the unsteady flow around a har-
monically oscillating airfoil, is to increase the lag
of the aerodynamic forces and moments behind the
deflection when the frequency is small. It is shown
theoretically that this will result in a serious deteri-
oration of the damping of both the lateral oscillation
and the high frequency longitudinal oscillation with
high Mach numbers. Use is made of derivatives
calculated for flutter purposes to estimate the un-
steady derivatives at aircraft oscillation frequencies.
Illustrative examples are presented.


N-34327*

Aeronautical Research Council (Gt. Brit.)
METHODS OF TESTING REINFORCED PLASTICS,
PARTS I AND I. F. T. Bartwell. 1954. 34p.
diagrs., photos., 13 tabs. (ARC R M 2702; ARC
11,375. Formerly ARC 10,095; Strut 1088)

An experimental comparison has been made between
five types of tensile tests including novel types de-
signed to enable axial loading conditions to be ap-
proached more readily than is the case with estab-
Ushed methods. Examination of the results of 240
tests indicates that significant differences can occur
between the results of different tests and that there
is also a significant variation between the properties
of material cut from different parts of the same
sheet. It Is concluded that the results, obtained when
testing paper-base material by novel methods, are
sufficiently good to justify development of a simpli-
fied apparatus of similar type for general use.


N-34328'

Aeronautical Research Council (Gt. Brit.)
THE INFLUENCE OF THICKNESS/CHORD RATIO
ON SUPERSONIC DERIVATIVES FOR OSCILLATING
AEROFOILS. W. P. Jones. 1954. 17p. diagrs.,
3 tabs. (ARC R& M 2679. Formerly ARC 10,871;
0.678; FM 1153)


5


By the use of Temple and Jahn's theory for the
oscillating flat plate and Busemann's theory for air-
foils in steady motion, derivatives are obtained for
symmetrical circular-arc and double-wedge airfoils
describing low frequency oscillations at supersonic
speeds. It is known that theoretically the torsional
aerodynamic damping for a flat plate oscillating
about an axis forward of the two-thirds chord posi-
tion is negative at low frequencies for a limited
range of supersonic speeds. In this report, however,
it is shown that the effect of increasing thickness/
chord ratio is to decrease the range of speeds for
which the aerodynamic damping is negative, and for
which the degree of freedom flutter is possible. The
present theory also allows for the forward movement
of the center of pressure from the half-chord position
as the airfoil thickness is increased, and leads to
better estimates of the stiffness derivatives for an
actual airfoil. In practice the center of pressure is
not at half-chord as predicted by linear theory.


N-34329'

Aeronautical Research Council (Gt. Brit.)
THEORETICAL CALCULATIONS OF THE DISTRI-
BUTION OF AERODYNAMIC LOADING ON A DELTA
WING. H. C. Garner. 1954. 33p. diagrs., 9 tabs.
(ARC R & M 2819. Formerly ARC 12,222; Perf. 538;
S & C 2295)

The distribution of velocity potential has been cal-
culated for a thin, flat, delta wing at small incidence.
The method introduces novel functions with 10
arbitrary constants to express the doublet distri-
bution over the wing and a special numerical integra-
tion to evaluate the downwash at 10 chosen points on
the surface. Three different forms of the doublet
distribution are employed and lead to three independ-
ent solutions of the resulting simultaneous equations.
One object of the paper is to form the first step
towards a fundamental comparison with measured
pressure distributions.


N-34331*

Forest Products Research Lab. (Gt.Brit.)
INVESTIGATION INTO GLUES AND GLUING.
PROGRESS REPORT EIGHTY-ONE JUNE 1954.
THE COMPARATIVE DURABILITY OF PLYWOOD
GLUES IN ENGLAND AND IN NIGERIA (SERIES
IV). FOURTH YEAR'S ANALYSIS. J. F. S.
Carruthers and M. S. Burridge. 10p., 3 tabs.
(Supersedes Progress Reports 62, 69, and 74)
(Forest Products Research Lab.)
Progress made on the evaluation of the comparative
durability of plywood adhesives in England and in
Nigeria is summarized. The adhesives evaluated
were of phenol-, resorchinol-, melamine-, and
fortified and normal urea-formaldehyde types. Test
results are presented for plywood joints bonded with
each type of adhesive over a 4-year period in four
different environments.


N-34383*

Royal Aircraft Establishment (Gt. Brit.)
EFFECT OF WING FLEXIBILITY ON LANDING
IMPACT OF A FLYING-BOAT HULL. D. Williams.
September 1954. 6p. (RAE Tech. Note Structures 136)





6


It is shown that wing flexibility cannot be the cause of
the discrepancy between theoretical and experimental
values for the interval between touchdown and max-
imum impact force.


N-34385*

Royal Aircraft Establishment (Gt.Brit.)
THE EFFECT OF HARD ANODIC COATINGS ON
THE FATIGUE STRENGTH OF D. T. D. 364B
ALUMINIUM ALLOY. E. G. Savage, E. G. F.
Sampson and J. K. Curran. June 1954. 10p.
diagrs., 4 tabs. (RAE Tech.Note Met. 200)

Wohler-type fatigue tests were made on D. T. D.
364B aluminum alloy test pieces which had been
anodised under conditions previously described in
order to produce hard coatings of about 0.001-or
0.002-inch thickness. Control tests were made on
unanodised test pieces. Film thickness, surface
roughness, and change of dimensions were also
determined. Hard anodising reduced the fatigue
strength of the test pieces by about one half, the
reductions being of the same order for each thick-
ness of coating. The anodic treatment increased the
average roughness of the surfaces from about 4 to
50 microinches and increased the diameter of the
test pieces by about 60 percent of the total film
thickness. A possible method of reducing the ill-
effects of hard anodising is described and further
work is proposed.


N-34386*

Royal Aircraft Establishment (Gt. Brit.)
SOME PROBLEMS IN AERODYNAMICS AND THEIR
SOLUTION BY ELECTRICAL ANALOGY. D.
Kuichemann and S. C. Redshaw. August 1954. 16p.
(RAE Tech.Note Aero 2323)
Some purely aerodynamical phenomena, which might
profitably be investigated by means of electrical
analogue computers, are described. No attempt has
been made to furnish a complete catalogue of prob-
lems but rather to present current issues, the solu-
tion of which would aid the development of practical
aerodynamics. Consideration has only been given to
those questions which can be solved more suitably by
an analogue computer rather than by a digital
machine, although the feasibility of partially solving
certain of them by the use of an analogue computer,
and completing the work with a digital machine, is
mentioned. In scope, the survey includes two- and
three-dimensional airfoils as well as interference
and a number of special problems. The relative
merits of the various forms of analogue computers
are not extensively discussed.


N-34388*

Royal Aircraft Establishment (Gt. Brit.)
THE EFFECT OF HEATING STEELS TO MODER-
ATELY ELEVATED TEMPERATURES. G. Meikle
and M. S. Binning. June 1954. 18p. diagrs.,
5 tabs. (RAE Tech. Note Met. 199)

A review has been made of available information on
the strength of structural steels up to temperatures
of the order of 3500 C. Experiments have been
made to determine the strength of aircraft steels
from 1000 C to 3500 C and the effect of heating at
selected temperatures for different periods was


NACA
RESEARCH ABSTRACTS NO. 75

determined. The percentage loss in strength and
Young's Modulus was found to be fairly constant for
all the ferritic steels at any particular temperature.
No evidence was obtained that the length of time of
heating, up to 1000 hours, had any additional
deleterious effect.


N-34390*

Royal Aircraft Establishment (Gt. Brit.)
THE STRENGTH OF ANNEALED AND HEAT-
TREATED GLASS. F. G. J. Brown and J. Ellis.
July 1954. 60p. diagrs., 7 tabs. (RAE
Structures 167)

This report develops statistical methods of choosing
allowable design stresses for annealed and heat-
treated glass. The results are easy to apply but
additional fundamental knowledge of some properties
of glass is needed before they can be used to the
best advantage. The report draws attention to these
gaps in existing knowledge and makes recommenda-
tions for further research.


N-34392*

Ministry of Supply (Gt. Brit.)
THE CREEP AND FATIGUE PROPERTIES OF
SOME WROUGHT COMPLEX ALUMINIUM
BRONZES. J. McKeown, D. N. Mends, E. S. Bale
and A. D. Michael. September 1954. 27p. diagrs.,
photos., 5 tabs. (MOS S & TM 9/54)

The work described in this report is primarily
concerned with the effect of aluminum content on the
creep and fatigue properties of complex aluminum
bronzes containing 5-percent each of iron and
nickel; the effect of heat treatment on such alloys
and the effect of the addition of small amounts of
high melting point elements on the creep properties
of alloys containing 9- and 10.5-percent aluminum
with 5-percent each of iron and nickel.



N-34393*

Royal Aircraft Establishment (Gt. Brit.)
THE FLOW FIELD OF WEDGE-SHAPED AND
CONICAL FLAME FRONTS AT LOW SPEEDS.
J. Weber. August 1954. 12p. diagrs. (RAE
Tech. Note Aero 2318)

The analytical representation of flame fronts by
source sheets is discussed and examples are given
for the streamlines and the pressure distribution on
them for the incompressible flow past two-
dimensional wedge-shaped and for conical flame
fronts.


N-34394*

Royal Aircraft Establishment (Gt. Brit.)
A CURVE FOLLOWER FOR DISCONTINUOUS
CURVES ON 35mm. FILM. O. H. Lange and G. J.
Herring. August 1954. 16p. diagrs. (RAE
Tech. Note GW 334)

A photo-electric system is described by means of
which the voltage analogue of the ordinates of the
curve of a single-valued function can be obtained to






NACA
RESEARCH ABSTRACTS NO. 75

an accuracy of approximately 2 percent of the
maximum ordinate. No restriction is placed on the
maximum slope of the curve, which is fed into the
machine as a trace on a film which may be of any
length. Circuits for the electronic components are
given and some applications are indicated.



N-34401*

Royal Aircraft Establishment (Gt. Brit.)
CONSIDERATION OF SOME NEW LOW-DENSITY
MATERIALS. G. A. Horridge. June 1954. 18p.
diagrs., photos. (RAE Tech. Note Chem. 1232)

It would be an advantage from the point of view of
structural efficiency to have a material that can be
moulded to form primary structures with a specific
gravity considerably less than existing reinforced
plastics. There is available at present a range of
low-density core materials which are used for
sandwich construction but a consideration of foamed
resins has not yet shown any promise that they can
be used as a basis for a low-density primary struc-
tural material.. This is because absolute strengths
are low and orientated reinforcement in the foam has
not improved the mechanical properties for competi-
tion with existing primary materials. An alternative
method to reduce the specific gravity of reinforced
plastics lies in the use of glass tubes instead of
fibers as reinforcement. Results of tests on such
materials have so far been promising and theoretical
considerations show both the limits of the effective-
ness of such reinforcement and the direction in which
to aim to produce a material of given properties.


N-34402'

Marine Aircraft Experimental Establishment
(Gt. Brit.) METHOD FOR REDUCING SEAPLANE
TAKE-OFF DISTANCES TO STANDARD CON-
DITIONS. J. A. Hamilton. April 1954. 44p.
diagrs.. photos., 10 tabs. (MAEE F/Res/245)

In this report are developed methods for the re-
duction of seaplane take-off distances to standard
conditions of weight, wind.and ambient temperature.
The expressions derived are applicable to the
waterborne run and to the airborne run up to the
50-foot height point. The methods maybe applied to
take-off with simulated engine failure. The theoret-
ical results are in good agreement with measure-
ments made on reciprocating engine seaplanes of
9000 lb and 78, 000 lb weight for winds up to 20
knots, weight changes of 20 percent, and a tempera-
ture range of 20 to 320 C. However, the general
application of the corrections should be limited to
temperature changes of less than 100 C, weight
changes of less than 10 percent, and wind changes of
less than 10 knots. These ranges should be adequate
for the majority of flight trials conducted in one
location.


N-34406*

Aeronautical Research Council (Gt. Brit.)
THE GROWTH OF A VORTEX IN TURBULENT
FLOW. H. B. Squire. March 18, 1954. 6p.
diagrs. (ARC 16.666; FM 2053)


7


The growth of a line vortex with time and the spread
of a trailing vortex behind a wing due to turbulence
are considered. It is shown that the eddy viscosity
for this type of motion may be taken to be propor-
tional to the circulation round the vortex and the
solution is then similar to the solution for the
growth of a vortex in laminar flow. The method is
applied to calculate the distance behind a wing for
which the trailing vortices will touch one another.









MISCELLANEOUS




NACA TN 1401

Errata No. 1 on "INTRODUCTION TO THE
PROBLEM OF ROCKET-POWERED AIRCRAFT
PERFORMANCE. H. Reese Ivey, Edward N.
Bowen, Jr. and Lester F. Oborny. December 1947.




NACA TN 3201

Errata No. 1 on "DIRECTIONAL STABILITY
CHARACTERISTICS OF TWO TYPES OF TANDEM
HELICOPTER FUSELAGE MODELS." James L.
Williams. May 1954.




NACA TN 3261

Errata No. 1 on "A METHOD FOR EVALUATING
THE EFFECTS OF DRAG AND INLET PRESSURE
RECOVERY ON PROPULSION-SYSTEM PERFORM-
ANCE.' Emil J. Kremzier. August 1954.



N-25292'

Advisory Group for Aeronautical Research and
Development. DESIGN AND OPERATION OF
INTERMITTENT SUPERSONIC WIND TUNNELS.
A. Ferri and S. M. Bogdonoff. May 1954. 108p.
diagrs., photo. (AGARDograph 1)

A study has been made of the operating character-
istics of intermittent wind tunnels, particularly of
the slowdown type. For the same test Mach and
Reynolds numbers, the blowdown tunnel is con-
siderably cheaper to build and to operate than the
comparable continuous tunnel. For tests of many
different setups, flexibility may be more important
than long running time. Typical construction, test
setup, instrumentation, and extensions to the tran-
sonic and hypersonic region have been discussed.
Some examples of operating intermittent tunnels are
also given.





8



UNPUBLISHED PAPERS



N-33708*

FLOW NEAR THE FORWARD TIP OF A HIGHLY
SWEPT WING AT MEDIUM ANGLES OF ATTACK -
PART I. (Ecoulement au Voisinage de la Pointe
Avant d'une Aile a Forte Fleche aux Incidences
Moyennes). Robert Legendre. (Presented at
Eighth International Congress of Theoretical and
Applied Mechanics, Istanbul, August 1952).
November 1954. 18p. diagrs. (Trans. from
Recherche A6ronautique, no. 30, 1952, p. 3-8)

An interpretation is made of the turbulent flow which
originates at the forward tip of a highly swept wing
as soon as the angle of attack is about 100. The case
considered is a plane, triangular, infinitely slender
wing. The method employed is Jones' method of
transverse sections. When the angle of attack of a
highly swept wing is sufficient, well-defined vortices
appear on the upper side of the wing. They cor-
respond approximately to the field which would be
induced by two rectilinear free vortices originating
at the wing tip.



N-33709*

FLOW NEAR THE FORWARD TIP OF A HIGHLY
SWEPT WING AT MEDIUM ANGLES OF ATTACK -
PART II. (Ecoulement au Voisinage de la Pointe
Avant d'une Aile a Forte Fleche aux Incidences
Moyennes). Robert Legendre. (Presented at
Eighth International Congress of Theoretical and
Applied Mechanics, Istanbul, August 1952).
November 1954. 12p. diagrs. (Trans. from
Recherche Aeronautique, no. 31, Jan. -Feb., 1953,
p.3-6)

The flow which originates in the neighborhood of the
forward tip of a highly swept wing at medium angles
of attack was previously schematized by the field of
two isolated vortices. This study replaces the
schematic representation by the field of the vortex
sheets attached to the two leading edges of the wing.
It is limited to the setting-up and the discussion of
the integral-differential equation of the problem.




DECLASSIFIED NACA REPORTS



THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL, 11/30/54:

RM E8B27a
RM E8B27b
RM E9B01
RM L52B29
RM L52C19
RM L52D28
RM L53B27
RM L53C27


NACA
RESEARCH ABSTRACTS NO. 75




NACA RM A50G06

LOW-SPEED INVESTIGATION OF A 0.16-SCALE
MODEL OF THE X-3 AIRPLANE LONGITUDINAL
CHARACTERISTICS. Noel K. Delany and Nora-Lee
F. Hayter. September 8, 1950. 80p. diagrs.,
photos., 2 tabs. (NACA RM A50G06) (Declassified
from Confidential, 10/29/54)

The low-speed, static, longitudinal characteristics of
a model of the X-3 airplane have been determined
with the wing flaps neutral and deflected. An inves-
tigation was also made of the effects of jettisonable-
nose fins, landing gear and doors, scoops, and pilot's
canopy on the longitudinal characteristics of the
model. The X-3 airplane utilizes a wing of aspect
ratio 3.01, a 4.5-percent-thick hexagonal section,
and a taper ratio of 0.4. The wing was equipped with
plain leading-edge flaps and split trailing-edge flaps.





NACA RM A50H02

EFFECT OF DIHEDRAL CHANGE ON THE THEO-
RETICAL DYNAMIC LATERAL RESPONSE CHAR-
ACTERISTICS OF A LOW-ASPECT-RATIO
STRAIGHT-WING SUPERSONIC AIRPLANE.
Donovan R. Heinle. December 7, 1950. 35p. diagrs.,
2 tabs. (NACA RM A50H02) (Declassified from
Confidential, 10/29/54)

Theoretical calculations were made of the lateral-
stability characteristics of a supersonic straight-
wing low-aspect-ratio airplane. A Reeves Electronic
Analog Computer was used to plot time histories of
the lateral response to various distrubances and to
obtain values of period, damping, and the roll excita-
tion parameter. The effect of a change to negative
geometric dihedral and changes in the yawing moment
due to roll derivative (Cn) were investigated.






NACA RM A51A16

LOW-SPEED INVESTIGATION OF A 0.16-SCALE
MODEL OF THE X-3 AIRPLANE LATERAL AND
DIRECTIONAL CHARACTERISTICS. Noel K. Delany
and Nora-Lee F. Hayter. March 16, 1951. 56p.
diagrs., photos., tab. (RM A51A16) (Declassified
from Confidential, 10/29/54)

A wind-tunnel investigation has been made of the low-
speed, static, lateral, and directional characteristics
of a model of an early design of the X-3 airplane with
the wing flaps neutral and deflected. Measurements
were also made of the fluctuations in rolling moment
with time. The model utilized a wing having an as-
pect ratio of 3.01, a 4.5-percent-thick hexagonal
section, and a taper ratio of 0.4. The wing was
equipped with plain leading-edge flaps and split trail-
ing edge flaps.





NACA
RESEARCH ABSTRACTS NO. 75



NACA RM A51H23

AERODYNAMIC CHARACTERISTICS OF A
LEADING-EDGE SLAT ON A 350 SWEPT-BACK
WING FOR MACH NUMBERS FROM 0.30 TO 0.88.
John A. Kelly and Nora-Lee F. Hayter. December
1951. 49p. diagrs., tab. (NACA RM A51H23)
(Declassified from Confidential, 11-30-54)
North American Aviation, Inc., has furnished the
NACA with data obtained from tests which were made
at Mach numbers from 0.300 to 0.883 to determine
the lift, drag, and pitching-moment characteristics
of a semispan model of a 350 sweptback wing with a
leading-edge slat. The wing was tested in the pres-
ence of a fuselage. Measurements were made of the
pressures acting on the slat in the retracted and full-
open positions. These pressure data were analyzed
to ascertain the slat opening-force characteristics
for two possible circular-arc slat tracks which could
be used for automatic operation of the slat.




NACA RM A51J25a

A SUMMARY OF AVAILABLE KNOWLEDGE CON-
CERNING SKIN FRICTION AND HEAT TRANSFER
AND ITS APPLICATION TO THE DESIGN OF HIGH-
SPEED MISSILES. Morris W. Rubesin, Charles B.
Rumsey and Steven A. Varga. November 1951. 17p.
diagrs. (NACA RM A51J25a) (Declassified from
Confidential, 10/29/54)

A review is made of the existing information con-
cerning boundary-layer characteristics: the temper-
ature recovery, the skin-friction coefficients, the
heat-transfer coefficients of both the laminar and
turbulent boundary layers, and the position of the
transition from laminar to turbulent flow. Compari-
son is made between existing flight data and results
computed by the boundary-layer momentum-integral
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.



NACA RM A52119a

INVESTIGATION OF AN NACA 4-(5)(05)-041 FOUR-
BLADE PROPELLER WITH SEVERAL SPINNERS AT
MACH NUMBERS UP TO 0.90. Robert M. Reynolds,
Donald A. Buell and John H. Walker. December
1952. 86p. diagrs., photos., 6 tabs. (NACA
RM A52119a) (Declassified from Confidential,
10,'29,'54)

This paper presents the propeller characteristics of
the NACA 4-(5)(05)-041 four-blade, single-rotation
propeller in combination with two NACA 1-series
spinners and two extended cylindrical spinners for a
wide range of blade angles at Mach numbers up to
0.90. Included are data for the propeller operating
at negative thrust at low forward speeds and operat-
ing at near-static conditions. A description of the
1000-horsepower propeller dynamometer is included.


9



NACA RM A52J17

CORRELATION OF BUFFET BOUNDARIES PRE-
DICTED FROM WIND-TUNNEL TESTS WITH THOSE
MEASURED DURING FLIGHT TESTS ON THE F8F-1
AND X-1 AIRPLANES TRANSONIC-BUMP METH-
OD. Andrew Martin and James F. Reed.
December 1952. 22p. diagrs., photos., tab.
(NACA RM A52J17) (Declassified from Confidential,
10/29/54)

Wing-fuselage-combination models of the F8F-1 and
X-1 airplanes have been tested in the Ames 16-foot
high-speed wind tunnel utilizing the transonic-bump
technique. Presented for these models are the lift
and pitching-moment coefficients, and fluctuating
wing-bending-moment and wake-pressure data.
Comparisons are presented of buffet boundaries
estimated from these data with flight-determined
buffet boundaries.



NACA RM E52A14

DYNAMIC RESPONSE OF TURBINE-BLADE
TEMPERATURE TO EXHAUST-GAS TEMPERA-
TURE FOR GAS-TURBINE ENGINES. Richard Hood
and William E. Phillips, Jr. February 1952. 41p.
photos., diagrs. (NACA RM E52A14) (Declassified
from Confidential, 11-30-54)

The frequency response of blade temperature to ex-
haust gas temperature is presented for two locations
in the blade and at several operating conditions. The
frequency response was determined by Fourier anal-
ysis of transient data. Two analytical methods are
presented, and the results are compared with experi-
mental data. The dynamic response of turbine-blade
temperature to exhaust-gas temperature exhibited
the form of an approximate first-order lag. The re-
sults are used to predict blade temperature for a
typical controlled and uncontrolled gas-turbine
engine.





10




NACA RM L8L31

THEORETICAL INVESTIGATION OF THE DYNAMIC
LATERAL STABILITY CHARACTERISTICS OF
DOUGLAS DESIGN NO. 39C, AN EARLY VERSION
OF THE X-3 RESEARCH AIRPLANE. Charles V.
Bennett. January 18, 1949. 39p. diagrs., 2 tabs.
(NACA RM L8L31) (Declassified from Confidential,
10/29/54)

Contains results of calculations made to determine
the neutral oscillatory stability boundaries, period
and time to damp of the oscillatory mode, and
motions following distrubances. The calculations
were made for Mach numbers of 0.75 and 2.3 at an
altitude of 35,000 feet and for the landing condition
at sea level.





NACA RM L9K08

STABILITY AND CONTROL CHARACTERISTICS AT
LOW SPEED OF A 1/4-SCALE BELL X-5 AIRPLANE
MODEL. LONGITUDINAL STABILITY AND CON-
TROL. William B. Kemp, Jr. Robert E. Becht and
Albert G. Few, Jr. March 14, 1950. 51p. diagrs.,
photos. (NACA RM L9K08) (Declassified from
Confidential, 10/29/54)

An investigation was conducted to determine the
longitudinal stability and control characteristics at
low speed of a 1/4-scale model of a preliminary X-5
airplane design. Tests were made through a range
of wing sweep angles from 200 to 600 and several
flap and slat configurations were investigated.






NACA RM L9L06

SPIN-TUNNEL INVESTIGATION OF A MODEL OF
A 600 DELTA-WING AIRPLANE TO DETERMINE
THE SPIN, RECOVERY, AND LONGITUDINAL
TRIM CHARACTERISTICS THROUGHOUT AN
EXTENSIVE RANGE OF MASS LOADINGS. Walter
J. Klinar and Ira P. Jones, Jr. February 15, 1950.
56p. diagrs., photos., 6 charts, 4 tabs.
(NACA RM L9L06) (Declassified from Confidential,
10/29/54)

An investigation has been conducted in the Langley
20-foot free-spinning tunnel to determine the spin
and recovery characteristics of a 600 delta-wing
model for a wide range of mass loadings. Varia-
tions were made in the relative density, center of
gravity, and inertia yawing-moment parameter. In
addition to the normal single-tail configuration,
dual-vertical-tail configurations were also investi-
gated. Glide tests were performed on the model
and force tests were conducted from 00 to 900 angle
of attack both on the model used for the dynamic
tests and on a somewhat larger model to determine
if any unusual trimming tendencies above the stall
existed.


NACA
RESEARCH ABSTRACTS NO. 75



NACA RM L50B28

THEORETICAL INVESTIGATION OF THE DYNAMIC
LATERAL STABILITY CHARACTERISTICS OF THE
DOUGLAS X-3 RESEARCH AIRPLANE, STUDY
41-B. Charles V. Bennett. April 27, 1950. 31p.
diagrs., 3 tabs. (NACA RM L50B28) (Declassified
from Confidential, 10/29/54)

Contains results of calculations made to determine
the period and damping of the lateral oscillation.
The calculations were made for a Mach number of
0.75 at sea level and 35,000 feet and for Mach num-
bers of 1.2 and 2.0 at 35,000 and 50,000 feet. The
effects of adding a rate gyro and the effects of
certain geometric modifications were studied.








NACA RM L50C17a

STABILITY AND CONTROL CHARACTERISTICS AT
LOW SPEED OF A 1/4-SCALE BELL X-5 AIRPLANE
MODEL. LATERAL AND DIRECTIONAL STABILITY
AND CONTROL. William B. Kemp, Jr. and Robert
E. Becht. June 20, 1950. 97p. diagrs., photos.
(NACA RM L50C17a) (Declassifiedfrom Confidential,
10/29/54)

An investigation was conducted to determine the lat-
eral and directional stability and control characteris-
tics at low speed of a 1/4-scale model of a prelimi-
nary Bell X-4 airplane design. Tests were made
through a range of sweep angles from 200 to 600, and
several flap and slat configurations were investigated.









NACA RM L50E02a

LONGITUDINAL CHARACTERISTICS AT MACH
NUMBER OF 1.24 OF 1/30-SCALE SEMISPAN
MODEL OF BELL X-5 VARIABLE-SWEEP AIR-
PLANE WITH WING SWEPT BACK 600 FROM TESTS
BY NACA WING-FLOW METHOD. Norman S.
Silsby, Garland J. Morris and Robert M. Kennedy.
October 12, 1950. 17p. diagrs., photos., tab.
(NACA RM L50E02a) (Declassified from Confidential,
10/29/54)

Tests were made by the NACA wing-flow method to
determine the longitudinal characteristics of a 1/30-
scale semispan model of the Bell X-5 variable-sweep
airplane with the wing in the 600 sweptback position
and with tail incidences of -2o and -6. Lift, drag,
and pitching-moment results for various angles of
attack are presented for a Mach number of 1.24.
The Reynolds number was about 1 x 106.





NACA
RESEARCH ABSTRACTS NO. 75



NACA RM L50F23

STABILITY CHARACTERISTICS AT LOW SPEED OF
A 1,/4-SCALE BELL X-5 AIRPLANE MODEL WITH
VARIOUS MODIFICATIONS TO THE BASIC MODEL
CONFIGURATIONS. Robert E. Becht and Albert G.
Few, Jr. August 16, 1950. 47p. diagrs., photo.
(NACA RM L50F23) (Declassified from Confidential,
10,'29, 54)

An investigation was conducted to determine the low-
speed longitudinal and lateral stability characteris-
tics of a 1/4-scale model of a preliminary Bell X-5
airplane design incorporating dive brakes and several
wing and fuselage modifications. Tests were made
at 200 and 600 wing-sweep angles.







NACA RM L50I28

THE EFFECT OF SWEEPBACK ON THE LONGITU-
DINAL CHARACTERISTICS AT A MACH NUMBER
OF 1.24 OF A 1/30-SCALE SEMISPAN MODEL OF
THE BELL X-5 AIRPLANE FROM TESTS BY THE
NACA WING-FLOW METHOD. Garland J. Morris,
Robert M. Kennedy and Norman S. Silsby.
November 27, 1950. 22p. diagrs., photos., tab.
(NACA RM L50128) (Declassified from Confidential,
10/29/54)

Tests were made by the NACA wing-flow method to
determine the effect of sweepback angle on the lon-
gitudinal characteristics of a 1/30-scale semispan
model of the Bell X-5 variable-sweep airplane with
wings in the 400 and 500 positions. A test was also
made with the fuselage alone. Lift, drag, and
pitching-moment results for various angles of attack
are presented for a Mach number of 1.24. The
results are compared with the results of a previous
test with the wing in the 600 position. The Reynolds
number was about 1 x 106.








NACA RM L50J27

STABILITY AND CONTROL CHARACTERISTICS OF
A 1/4-SCALE BELL X-5 AIRPLANE MODEL IN THE
LANDING CONFIGURATION. Robert E. Becht. 38p.
diagrs., photos. (NACA RM L50J27) (Declassified
from Confidential, 10/29/54)

An Investigation was conducted to determine the sta-
bility and control characteristics of a 1/4-scale
model of a preliminary Bell X-5 airplane design in
the landing configuration with and without dive brakes.
Tests were made at wing sweep angles of 200and 609


11




NACA RM L50K15

LONGITUDINAL-CONTROL EFFECTIVENESS AND
DOWNWASH CHARACTERISTICS AT A MACH NUM-
BER OF 1.24 OF A 1/30-SCALE SEMISPAN MODEL
OF THE BELL X-5 AIRPLANE AS DETERMINED BY
THE NACA WING-FLOW METHOD. Richard H.
Sawyer, Robert M. Kennedy and Garland J. Morris.
January 8, 1951. 37p. diagrs., photos., tab.
(NACA RM L50K15) (Declassified from Confidential,
10/29/54)
An investigation was made at a Mach number of 1.24
by the NACA wing-flow method to determine the lon-
gitudinal control effectiveness and downwash charac-
teristics of a 1/30-scale semispan model of the vari-
able sweep Bell X-5 airplane with the wing of the
model swept back 600, 500, and 400. Lift, drag, and
pitching moments were obtained for various angles of
attack for several horizontal tail settings and with
tail off for each angle of sweep of the wing. The
Reynolds number was about 1 x 106.




NACA RM L50Llla

EFFECTS OF A FUSELAGE FLAP DIVE BRAKE ON
THE AERODYNAMIC CHARACTERISTICS OF 1/30-
SCALE SEMISPAN MODEL OF THE BELL X-5
VARIABLE-SWEEP AIRPLANE AT A MACH NUM-
BER 1.24 AS DETERMINED BY THE NACA WING-
FLOW METHOD. Robert M. Kennedy. February 8,
1951. 15p. diagrs., photos., tab. (NACA
RM L50Llla) (Declassified from Confidential,
10/29/54)

An investigation was made by the NACA wing-flow
method to determine the effect of a fuselage flap dive
brake on the longitudinal characteristics of a 1/30-
scale semispan model of the Bell X-5 variable-sweep
airplane with the wing in the 600 sweptback position.
Lift, drag, and pitching-moment results for various
angles of attack are presented for a Mach number of
1.24. The results are compared with the results of
a previous test without the dive brake. The Reynolds
number of the tests was about 1 x 106.




NACA RM L50L19

STATIC LONGITUDINAL STABILITY AND
DYNAMIC CHARACTERISTICS AT HIGH ANGLES
OF ATTACK AND AT LOW REYNOLDS NUMBERS
OF A MODEL OF THE X-3 SUPERSONIC RESEARCH
AIRPLANE. Sanger M. Burk, Jr. and Burton E.
Hultz. February 6, 1951. 76p. diagrs., photos.,
4 tabs. (NACA RM L50L19) (Declassified from
Confidential, 10/29/54)

An investigation was conducted in the Langley
20-foot free-spinning tunnel on a dynamic model
and a static model representative of a supersonic
airplane having a long nose and small wing to
determine the motions and trim conditions possible










at high angles of attack. Tests were also conducted
in the Langley 300 mph 7- by 10-foot tunnel on the
static model to determine the effect of higher
Reynolds numbers. The results of the investigation,
in general, indicate that scale effect may appreciably
influence free dynamic results obtained at low
Reynolds numbers and high angles of attack on a
design having a very long nose and a very small
wing.




NACA RM L51H20a

SOME FLYING-QUALITIES STUDIES OF A TANDEM
HELICOPTER., Kenneth B. Amer. October 1951.
29p. diagrs., photos. (NACA RM L51H20a)
(Declassified from Confidential, 11-30-54)

Initial results of an investigation of the flying qual-
ities of a representative tandem helicopter are pre-
sented. A comparison is made with previously pub-
lished flying-qualities requirements based on studies
of several single-rotor helicopters. A method is
presented which gives promise of improving the lon-
gitudinal flying qualities of tandem helicopters with
little weight penalty.




NACA RM L51I25

PRESSURE DISTRIBUTION AT LOW SPEED ON A
1/4-SCALE BELL X-5 AIRPLANE MODEL.
William B. Kemp, Jr. and Albert G. Few, Jr.
December 1951. 86p. diagrs., photos. (NACA
RM L51I25) (Declassified from Confidential,
10/29/54)

An investigation was conducted at 200 and 600 sweep
to determine the low-speed pressure distribution on
the wing, wing slat, wing leading-edge fillet, and
fuselage of a 1/4-scale model of a preliminary Bell
X-5 airplane design. Included in the tests were in-
vestigations of nose-inlet velocity ratios with the jet
exit duct fully open and partially closed.




NACA RM L51J18

WIND-TUNNEL INVESTIGATION OF THE EFFECTS
OF VARIOUS ASYMMETRIC CANOPY MODIFICA-
TIONS ON THE BEHAVIOR OF DESCENDING PARA-
CHUTES. Stanley H. Scher. February 1952. 20p.
diagrs., photos., tab. (NACA RM L51J18)
(Declassified from Confidential, 11/1/54)

Presented herein are the results of free-spinning
tunnel tests made to determine the behavior in de-
scent of eighteen 14-inch-diameter hemispherical
and quasi-conical parachutes, having some various
types of asymmetric canopy modifications and each
having a cylindrical weight suspended below the para-
chute. The asymmetric modifications were made in
an attempt to provide an appreciable lateral flight
component of the parachute and cylinder.


NACA
RESEARCH ABSTRACTS NO. 75



NACA RM L51K22

INVESTIGATION OF THE LOW-SPEED AERODY-
NAMIC CHARACTERISTICS OF A VARIABLE-
SWEEP AIRPLANE MODEL WITH A TWISTED AND
CAMBERED WING. William B. Kemp, Jr., Robert
E. Becht and Albert G. Few, Jr. February 1952.
62p. diagrs., photos. (NACA RM L51K22)
(Declassified from Confidential, 10/29/54)

An investigation was conducted to determine the
aerodynamic characteristics at low speed of a
variable sweep airplane model with a twisted and
cambered wing designed to have a uniform load
distribution at a Mach number of 1.10 and a lift
coefficient of 0.25 and 500 sweep. Comparison is
made with data obtained on the same model with an
untwisted, uncambered wing of the same geometric
plan form. Also included in the investigation is the
effect of partial-span split flaps.





NACA RM L51K27

INVESTIGATION OF THE DISTRIBUTION OF LIFT,
DRAG, AND PITCHING MOMENT BETWEEN THE
WING AND FUSELAGE OF A 1/30-SCALE SEMI-
SPAN MODEL OF THE BELL X-5 AIRPLANE AT A
MACH NUMBER OF 1.24 BY THE NACA WING-
FLOW METHOD. Norman S. Silsby and Garland J.
Morris. January 1952. 32p. diagrs., photos.,
tab. (NACA RM L51K27) (Declassified from
Confidential, 10/29/54)

An investigation was made at a Mach number of 1.24
by the NACA wing-flow method to determine the
distribution of lift, drag, and pitching moment be-
tween the wing and fuselage, the root bending mo-
ments of the wings, and the effect of wing flexibility
(at 600 sweep) on the longitudinal stability charac-
teristics of a 1/30-scale semispan model of the
variable-sweep Bell X-5 airplane. Lift, drag,
pitching moments, and wing root bending moments
were obtained for various angles of attack. The
Reynolds number was about 1.0 x 106.





NACA RM L52G21

THE VERTICAL-TAIL LOADS MEASURED DURING
A FLIGHT INVESTIGATION ON A JET-POWERED
BOMBER AIRPLANE. T. V. Cooney. May 1953.
32p. diagrs., photo. (NACA RM L52G21) (Declas-
sified from Confidential, 11-30-54)

The vertical-tail loads measured in rudder kicks and
gradual sideslips during a flight investigation of a jet
bomber are presented, and the results are given of
an analysis of the measured data to determine the
vertical-tail lift-coefficient parameters CL, and
CL5. The flight tests were performed at two
altitudes, 20,000 feet and 30,000 feet, and covered
the Mach number range from 0.30 to 0.76.





NACA
RESEARCH ABSTRACTS NO.75



NACA RM L52119a

THE EFFECTS OF A SMALL JET OF AIR EX-
HAUSTING FROM THE NOSE OF A BODY OF REV-
OLUTION IN SUPERSONIC FLOW. Eugene S. Love.
November 1952. 45p. diagrs., photos. (NACA
RM L52I19a) (Declassified from Confidential,
10/29/54)

Results are presented which show the effects at a
Mach number of 1.62 of a small jet of air exhausting
from the nose of an elliptical body of revolution upon
boundary-layer transition and the viscous, pressure,
and total drag of the forebody. Tests were conduct-
ed at Reynolds numbers of 2.13 x 106 and
7.66 x 106. The maximum range of thrust coef-
ficients for the small jet was from 0 to about -0.085.




NACA RM L52122

AN ANALYSIS OF ESTIMATED AND EXPERIMENT-
AL TRANSONIC DOWNWASH CHARACTERISTICS
AS AFFECTED BY PLAN FORM AND THICKNESS
FOR WING AND WING-FUSELAGE CONFIGURA-
TIONS. Joseph Weil, George S. Campbell and
Margaret S. Diederich. November 1952. 92p.
diagrs., 2 tabs. (NACA RM L52I22) (Declassified
from Confidential, 10/29/54)

This paper presents a summary of the effects of
changes in wing plan form and thickness ratio on
the downwash characteristics of wing and wing-
fuselage configurations in the Mach number range
between 0. 6 and 1. 1. Data obtained by the
transonic-bump technique at two tail heights have
been compared with theoretical estimations made in
the subsonic and supersonic Mach number range.





NACA RM L52J09

INVESTIGATION AT TRANSONIC SPEEDS OF THE
EFFECT OF A POSITIVE-LIFT BALANCING TAB
ON THE HINGE-MOMENT AND LIFT CHARACTER-
ISTICS OF A FULL-SPAN FLAP ON A TAPERED
450 SWEPTBACK WING OF ASPECT RATIO 3.
Vernard E. Lockwood and Joseph E. Fikes.
November 1952. 22p. diagrs., photo. (NACA
RM L52J09) (Declassified from Confidential,
10/29/54)

Lift and hinge moments of a wing with a flap and
balancing tab were obtained at transonic speeds by
testing in the high-velocity field over a reflection
plane on the side wall of the Langley high-speed 7-
by 10-foot tunnel. The wing has a quarter-chord-
line sweepback of 45.580 and aspect ratio of 3, and
a taper ratio of 0.5 and was approximately 7.6 per-
cent thick. The control consisted of 25.4-percent-
chord full-span flap linked to an auxiliary lifting
surface. The investigation was made over a
limited tab-deflection and angle-of-attack range
from M = 0.70 to 1.10 at a flap deflection of 3.9.


13




NACA RM L52J10

EFFECTS OF SWEEPBACK AND TAPER ON THE
FORCE AND CAVITATION CHARACTERISTICS OF
ASPECT-RATIO-4 HYDROFOILS. Douglas A.
King and Norman S. Land. December 1952. 61p.
diagrs., photos., tab. (NACA RM L52J10)
(Declassified from Confidential, 10/29/54)

An investigation was made to determine the effect
of sweep and taper ratio on the cavitation and
hydrodynamic characteristics of semispan hydrofoils
of aspect ratio 4 and NACA 65A006 airfoil section.
Sweep increased cavitation speed slightly, and
taper had negligible effect. The large hydrodynamic
twisting of the 600 swept hydrofoil obscured the
effects of sweep. Complete cavitation limited the
lift attainable to about 2000 lb/sq ft, 1210 lb/sq ft
for the hydrofoil with 600 sweep. Chordwise pro-
gression of cavitation on unswept hydrofoil moved
the center of pressure rearward.






NACA RM L52J13

A WIND-TUNNEL INVESTIGATION AT LOW
SPEEDS OF THE AERODYNAMIC CHARACTER-
ISTICS OF VARIOUS SPOILER CONFIGURATIONS
ON A THIN 600 DELTA WING. Harleth G. Wiley
and Martin Solomon. November 1952. 20p. diagrs.
(NACA RM L52J13) (Declassified from Confidential,
10/29/54)

Results are presented of a preliminary wind-tunnel
investigation at low speeds and angles of attack up
to 360 to determine the rolling- and yawing-
moment characteristics of various locations and
configurations of spoilers on a thin 600 delta wing of
constant thickness. Presented also are the effects
of wing slots and spoiler perforations.







NACA RM L52J24

CHARACTERISTICS OF FLAP-TYPE SPOILER
AILERONS AT VARIOUS LOCATIONS ON A 600
DELTA WING WITH A DOUBLE SLOTTED FLAP.
Delwin R. Croom. December 1952. 31p. diagrs.,
3 tabs. (NACA RM L52J24) (Declassified from
Confidential, 10/29/54)

This paper contains results and discussion of a wind-
tunnel investigation at low speeds of spoiler-type
controls located at various chordwise and spanwise
positions on a thin delta wing equipped with a double-
slotted flap. Results indicated a spoiler located on
the flap gave rolling-moment coefficients that varied
fairly linearly with spoiler projections and were
about the same magnitude for the flap-retracted or
the flap-deflected condition.






14





NACA RM L52J31
National Advisory Committee for Aeronautics.
MEASUREMENTS OF FLUCTUATING PRESSURES
ON A 1/4-SCALE MODEL OF THE X-1 AIRPLANE
WITH A 10-PERCENT-THICK WING IN THE
LANGLEY 16-FOOT TRANSONIC TUNNEL.
Louis W. Habel and Seymour Steinberg. January
1953. 29p. diagrs., photos. (NACA RM L52J31)
(Declassified from Confidential, 11-30-54)
Pressure fluctuations have been measured near the
trailing edge of the wing and near the leading edge of
the tail of a 1/4-scale mode of the X-1 airplane with
a 10-percent-thick wing in the Langley 16-foot
transonic tunnel. Results are presented for a Mach
number range from 0.70 to 1.00 and an angle-of-
attack range from about -40 to about 150 with and
without vortex generators installed on the wings.
Frequency analyses of the pressure fluctuations
made while tests were in progress are presented
for several test conditions.


NACA
RESEARCH ABSTRACTS NO. 75




NACA RM L52K19
National Advisory Committee for Aeronautics.
AERODYNAMIC CHARACTERISTICS OF TWO
DELTA WINGS AT MACH NUMBER 4.04 AND COR-
RELATIONS OF LIFT AND MINIMUM-DRAG DATA
FOR DELTA WINGS AT MACH NUMBERS FROM 1.62
TO 6.9. Edward F. Ulmann and Robert W. Dunning.
December 1952. 20p. diagrs. (NACA RM L52K19)
(Declassified from Confidential, 10/29/54)

Tests were made in the Langley 9- by 9-inch Mach
number 4 blowdown jet to obtain the aerodynamic
characteristics of two semispan delta wings of aspect
ratio 2.31 having 8-percent-thick double-wedge air-
foil sections and maximum-thickness locations at 18
and 60 percent chord. The data are presented to-
gether with the results of previous tests on the same
and similar delta wings at other Mach numbers and
show methods of correlating lift and drag data on
sharp-edge delta wings at supersonic Mach numbers.


SNACA-Langley 12-21-54 4M






NACA



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