Research abstracts

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Research abstracts
Physical Description:
93 v. : ; 27 cm.
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English
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United States -- National Advisory Committee for Aeronautics
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National Advisory Committee for Aeronautics
Place of Publication:
Washington, D.C
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irregular
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Subjects / Keywords:
Aeronautics -- Abstracts -- Periodicals   ( lcsh )
Aeronautics -- Research -- Abstracts -- Periodicals   ( lcsh )
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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).

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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:
AA00009235:00069

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National Advisory Committee for Aeroi



Research Abstracts


OCTOBER 1,


CURRENT NACA REPORTS


NACA Rept. 997
National Advisory Committee for Aeronautics.
SUMMARY OF INFORMATION RELATING TO GUST
LOADS ON AIRPLANES. Philip Donely. 1950.
ill, 51p. diagrs., photos., 21 tabs. (NACA Rept. 997.
Formerly TN 1976)
*
Available information on gust structure, airplane
reactions, and pertinent operating statistics has
been examined. This report attempts to coordinate
this information with reference to the prediction of
gust loads on airplanes. The material covered
represents research up to October 1947.


NACA Rept. 1000
National Advisory Committee for Aeronautics.
CALCULATION OF THE AERODYNAMIC LOADING
OF SWEPT AND UNSWEPT FLEXIBLE WINGS OF
ARBITRARY STIFFNESS. Franklin W. Diederich.
1950. ii, 29p. diagrs., 10 tabs. (NACA Rept. 1000.
Formerly RM LBG27a, TN 1876)

A method is presented for calculating the aero-
dynamic loading, the divergence speed, and certain
stability derivatives of swept and unswept wings and
tail surfaces of arbitrary stiffness. Provision is
made for using either stiffness curves and root rota-
tion constants or structural influence coefficients in
the analysis. Computing forms, tables of numerical
constants required in the analysis, and an illustrative
example are included to facilitate calculations by
means of the method.


NACA Rept. 1009
National Advisory Committee for Aeronautics.
INVESTIGATION OF FRETTING BY MICROSCOPIC
OBSERVATION. Douglas Godfrey. 1951. ii, 10p.
photos. (NACA Rept. 1009. Formerly TN 2039)

An experimental investigation, using microscopic ob-
servation and color motion photomicrographs of the
action, was conducted to determine the cause of fret-
ting. Glass and other noncorrosive materials, as
well as metals, were used as specimens. A very
simple apparatus vibrated convex surfaces in contact
with stationary flat surfaces. This led to the con-
clusion that fretting was initiated by the loosening,
due to inherent adhesive forces, of finely divided and
apparently virgin material, and that its initiation is
independent of vibratory motion or high sliding
speeds. The fretting of platinum, glass, quartz, ruby,
and mica relegated the role of oxidation as a cause to
that of a secondary factor. Fretting occurred to
clean nonmetals and metals readily, and glass


microscope slides and steel balls provided an excel-
lent method for visual studies.


NACA Rept. 1011
National Advisory Committee for Aeronautics.
DYNAMICS OF A TURBOJET ENGINE CON-
SIDERED AS A QUASI-STATIC SYSTEM. Edward W.
Otto and Burt L. Taylor, Il. 1951. ii, 12p. diagrs.
(NACA Rept. 1011. Formerly TN 2091)

A determination of the dynamic characteristics of a
typical turbojet engine with a centrifugal compres-
sor, a sonic-flow turbine-nozzle diaphragm, and
fixed-area exhaust nozzle is presented. A general-
ized equation for the transient behavior of the engine
was developed; this equation was then verified by
calculations using compressor- and turbine-
performance charts extrapolated from equilibrium
operating data and by experimental data obtained
from an engine operated under transients in fuel
flow. The results indicate that a linear differential
equation for engine acceleration as a function of fuel
flow and engine speed for operation near a steady-
state operating condition can be written.


NACA TN 2412
National Advisory Committee for Aeronautics.
THEORETICAL FORCE AND MOMENTS DUE TO
SIDESLIP OF A NUMBER OF VERTICAL TAIL
CONFIGURATIONS AT SUPERSONIC SPEEDS.
John C. Martin and Frank S. Malvestuto, Jr.
September 1951. 60p. diagrs., photos. (NACA TN
2412)

Formulas are obtained by means of the linearized
theory for the lateral force due to sideslip, yawing
moment due to sideslip, and the rolling moment due
to sideslip for norma I tail arrangements consisting
of rectangular, triangular, and sweptback vertical
tails of arbitrary taper and sweep mounted symmet-
rically on a horizontal tail of arbitrary shape. The
resitlts are restricted to cases where the leading
edges are supersonic. The effect of the horizontal
tail on the derivatives is evaluated for certain cases.
A series of design curves is presented for the de-
rivatives considered for certain taper ratios.


NACA TN 2429
National Advisory Committee for Aeronautics.
STUDY OF VORTEX SHEDDING AS RELATED TO
SELF-EXCITED TORSIONAL OSCILLATIONS OF
AN AIRFOIL. Raymond L. Chuan and Richard J.
Magnus, California Institute of Technology.
September 1951. 49p. diagrs. (NACA TN 2429)

Results of the experimental investigation of the self-
excited torsional oscillation of an NACA 0006 airfoil


NO. 8


#AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASH I NGTON 25, D. C., CITING CODE NUMBER, TITLE AND AUTHOR.





NACA
RESEARCH ABSTRACTS NO.8


suspended elastically are covered. The relationship
between torsional oscillation and shedding of
vortices was investigated for this airfoil. Two types
of oscillation phenomena were found in the investiga-
tion. One type, exhibited by cases with angles of
attack just above stall, persisted with increasing
velocity without reaching any apparent limit within
the range of velocity attainable in the present wind
tunnel. The other type, exhibited by cases with
higher angles of attack, only showed self-excited
oscillations in a certain range of velocity, the
range decreasing with increasing angle of attack.


NACA TN 2439
National Advisory Committee for Aeronautics.
A THEORY OF CONDUCTIVITY OF COLD-WORKED
COPPER. Rolf Landauer. September 1951. 23p.
diagrs. (NACA TN 2439)

The increase in the resistivity of copper under cold
working is calculated. The increase is assumed to
be caused by dislocations surrounded by a long-range
electrostatic field that scatters the conduction
electrons. The amount of scattering is found by the
method of deformation potentials of Bardeen and
Shockley. The scattering is present in addition to
the normal thermal scattering and is regarded as a
perturbation in the Boltzmann equation. This per-
turbation is used to find the incremental resistance
per dislocation. From this calculated increment in
resistance and the known increase of resistivity of
heavily cold-worked copper, the number of disloca-
tions in the cold-worked copper is found to be
5 x 1011/square centimeter, in agreement with

measurements.



NACA TN 2446
National Advisory Committee for Aeronautics.
WIDTH OF DEBYE-SCHERRER LINES FOR
FINITE SPECTRAL WIDTH OF PRIMARY BEAM.
Hans Ekstein, Armour Research Foundation.
September 1951. 9p. diagrs. (NACA TN 2446)

Width of a Debye-Scherrer line under the combined
influence of spectral width of the primary radiation
and of the small size of the crystal grains is cal-
culated, omitting all geometric causes of line
broadening. The shape of the crystals is assumed
spherical, with the same diameter for all grains.
The orientation is assumed to be random, with neg-
ligible statistical fluctuations. It is shown that the
width under these circumstances is the sum of that
width which is obtained with very large crystals
("spectral" width) and that which is obtained with
monochromatic radiation ("size" width).



NACA TN 2448
National Advisory Committee for Aeronautics.
X-RAY DIFFRACTION BY BENT CRYSTAL
LAMELLAE. Hans Ekstein, Armour Research
Foundation. September 1951. 14p. diagrs. (NACA
TN 2448)

A bent crystal lamella may be assumed to consist of
individual irregularly shaped blocks having mutual
angular disorientation or to be regular in structure
except for plastic deformation. The blocks were


assumed to be of such a size that the kinematic
theory would be valid. X-ray integrated intensity
was calculated in the first case by adding the inten-
sities of the individual blocks and in the second case
by adding the amplitudes. Theoretical analysis
indicated that, in both cases, moderate bending is
not expected to change the integrated intensity or to
cause broadening of the Debye-Scherrer lines.


NACA TN 2453
National Advisory Committee for Aeronautics.
AN EXPERIMENTAL STUDY OF WATER-
PRESSURE DISTRIBUTIONS DURING LANDINGS
AND PLANING OF A HEAVILY LOADED
RECTANGULAR FLAT-PLATE MODEL. Robert F.
Smiley. September 1951. 40p. dlagrs., 3 tabs.
(NACA TN 2453) 0

Water-pressure, velocity, draft, wetted length, and
acceleration measurements are presented for
smooth-water landing and planing tests of a rectan-
gular flat-plate model. Landings were made at
fixed trims between 60 and 45, for flight-path angles
between 2 and 20, with beam-loading coefficients
of 18.8 and 36.5. Planing runs were made for trims
between 60 and 45.0 For landings, the experimental
pressure coefficients based on the equivalent planing
velocity appear to be substantially independent of the
deceleration of the model. The peak pressures were
approximately equal to the dynamic pressure
corresponding to the velocity of the peak-pressure
point.


NACA TN 2457
National Advisory Committee for Aeronautics.
AIR FORCES AND MOMENTS ON TRIANGULAR
AND RELATED WINGS WITH SUBSONIC LEADING
EDGES OSCILLATING IN SUPERSONIC POTENTIAL
FLOW. Charles E. Watkins. September 1951. 44p.
diagrs. (NACA TN 2457)

This analysis treats the air forces and moments in
supersonic potential flow on oscillating triangular
wings and a series of sweptback and arrow wings
with subsonic leading edges and supersonic trailing
edges. The linearized velocity potential for the
wings undergoing sinusoidal torsional oscillations
simultaneously with vertical translations is derived
in the form of a power series in terms of a frequency
parameter. Although as many terms of such a
series expansion as may be desired can be deter-
mined, the terms after the first few become very
cumbersome. Closed expressions that include the
reduced frequency to the third power, an order
which is sufficient for a large class of practical ap-
plications, are given for the velocity potential and
components of chordwise section force and moment
coefficients.



NACA TN 2462
National Advisory Committee for Aeronautics.
INFLUENCE OF REFRACTION ON THE
APPLICABILITY OF THE ZEHNDER-MACH
INTERFEROMETER TO STUDIES OF COOLED
BOUNDARY LAYERS. Martin R. Kinsler.
September 1951. 39p. diagrs., tab. (NACA TN
2462)






NACA
RESEARCH ABSTRACTS NO.8 3


An analytical investigation was conducted in order to
determine the influence of light refraction on the
applicability of the Zehnder-Mach optical interfer-
ometer to two-dimensional cooled-boundary-layer
studies. A method is presented for estimating the
effects of refraction in a cooled boundary layer.
Some numerical results are presented for laminar
and turbulent boundary layers over a flat plate having
wall to free-stream temperature ratios from approxi-
mately 0.33 to 0.9, a free-stream Mach number of
1.0, and a Reynolds number of 5 x 10".


NACA TN 2463
National Advisory Committee for Aeronautics.
EXPERIMENTAL INVESTIGATION OF THE PRES-
SURE DISTRIBUTION ABOUT A YAWED CIRCULAR
CYLINDER IN THE CRITICAL REYNOLDS NUMBER
RANGE. William J. Bursnall and Laurence K.
Loftin, Jr. September 1951. 34p. diagrs. (NACA
TN 2463)

An experimental investigation was made of the pres-
sure distribution about a circular cylinder at var-
ious angles of yaw. The results indicate that the
flow and force characteristics in the range of
Reynolds number based on normal velocity compo-
nent near and above the critical cannot be determined
only by the component of flow normal to the cylinder
axis. Localized regions of laminar separation which
were present in the 0 yaw case in the supercritical
Reynolds number range became less distinct as the
yaw angle was increased and completely disappeared
for yaw angles of 45 and 600.


NACA TN 2464
National Advisory Committee for Aeronautics.
TWO AXIAL-SYMMETRY SOLUTIONS FOR
INCOMPRESSIBLE FLOW THROUGH A CENTRIFU-
GAL COMPRESSOR WITH AND WITHOUT INDUCER
VANES. Gaylord 0. Ellis, John D. Stanitz and
Leonard J. Sheldrake. September 1951. 34p. diagrs.
(NACA TN 2464)

Solutions for axially symmetric flow through an
impeller with and without inducer vanes were ob-
tained by relaxation methods using an analysis de-
veloped in the report. The fluid was considered to
be inviscid and incompressible. The impeller select-
ed was of arbitrary design except that the impeller
blades were assumed to have zero thickness and to
consist of radial elements. Plots of streamlines,
lines of constant velocity, and lines of constant
tangential blade force are presented and discussed.


NACA TN 2470
National Advisory Committee for Aeronautics.
EFFECT OF AN AUTOPILOT SENSITIVE TO
YAWING VELOCITY ON THE LATERAL STABILITY
OF A TYPICAL HIGH-SPEED AIRPLANE.
Ordway B. Gates, Jr. and Leonard Sternfield.
September 1951. 29p. diagrs., 2 tabs. (NACA TN
2470. Formerly RM L50F22)

Calculations have been made to determine the effect
of an autopilot sensitive to yawing velocity on the
lateral stability of a fighter airplane designed for
transonic or supersonic speeds. The effects of
inclination of the gyro reference axis and of time lag
in the autopilot are discussed. Airplane motions in


sideslip subsequent to a disturbance in sideslip are
also presented for several representative flight con-
ditions.


NACA TN 2477
National Advisory Committee for Aeronautics.
INVESTIGATION OF THE AIR-COMPRESSION PRO-
CESS DURING DROP TESTS OF AN OLEO-
PNEUMATIC LANDING GEAR. James H. Walls.
September 1951. 17p. diagrs., photo. (NACA TN
2477)

A brief study has been made to evaluate the impor-
tance of the type of air-compression process on the
loads produced on an oleo-pneumatic landing gear
during impact and to determine the type of air com-
pression process actually obtained during drop tests.
A simplified analysis to determine the effect which
different air-compression processes might have in-
dicates that the value of the air-compression expo-
nent should have relatively little effect on the landing
gear loads throughout most of the impact. The anal-
ysis of experimental data obtained in these tests
shows that the polytropic exponent ranged from 1.01
to 1.10 for the conditions tested. The general
trend of the data appears relatively independent of
vertical contact velocity.


NACA TN 2478
National Advisory Committee for Aeronautics.
A PROCEDURE FOR CALCULATING THE
DEVELOPMENT OF TURBULENT BOUNDARY
LAYERS UNDER THE INFLUENCE OF ADVERSE
PRESSURE GRADIENTS. Kennedy F. Rubert and
Jerome Persh. September 1951. 61p. diagrs.
(NACA TN 2478)

A procedure based on the kinetic-energy equation
and an extended form of the momentum equation has
been devised for calculating the development of
turbulent boundary layers in adverse pressure
gradients. Predictions, by this method, of turbulent-
boundary-layer development in comparison with
experimental results from several sources are pre-
sented for a number of cases of flow over flat plates
and airfoils and in conical diffusers. In the range of
boundary-layer flow short of separation, the agree-
ment with experiment is, in most cases, quite satis-
factory; in some instances, however, there

experimental results. It is believed, however, that
good agreement has been obtained in enough
instances to justify continuation of effort along the
present lines, particularly with respect to improve-
ment of the correlations and refinement of the
equations.


NACA TN 2479
National Advisory Committee for Aeronautics.
TABLES OF EXACT LAMINAR-BOUNDARY-LAYER
SOLUTIONS WHEN TH E WALL IS POROUS AND
FLUID PROPERTIES ARE VARIABLE. W. Byron
Brown and Patrick L. Donoughe. September 1951.
68p. diagrs., 2 tabs. (NACA TN 2479)

Exact solutions of the laminar boundary equations
were computed and tabulated for a range of fixed
values of Euler number, temperature ratio, and flow
through a porous wall. Euler numbers are 0, 0.5, 1,





NACA
RESEARCH ABSTRACTS NO.8


and negative values to the separation point. Tem-
perature ratios are 1, 2, and 4 for the impermeable
wall and for two values of coolant flow. In addition,
results from temperature ratios of 1/2 and 1/4 are
given for the impermeable wall. For each case,
boundary-layer thicknesses and heat-transfer and
friction coefficients were computed and tabulated.


NACA TN 2480
National Advisory Committee for Aeronautics.
COMPARISON OF HEAT TRANSFER FROM AIRFOIL
IN NATURAL AND SIMULATED ICING CONDITIONS.
Thomas F. Gelder and James P. Lewis. September
1951. 51p. diagrs., photos., 2 tabs. (NACA TN 2480)

An experimental investigation of the heat transfer
from an 8-foot-chord airfoil model in clear air and
in simulated icing conditions was conducted in the
icing tunnel. These results are compared with those
obtained in a flight investigation with the same model
at similar operating conditions. The tunnel results
indicate the effect of tunnel turbulence by the forward
movement of transition from laminar to turbulent
heat transfer. The flight results indicate that the
convective heat transfer in icing is considerably
different from that measured in clear air and only
slightly different from that obtained in the tunnel
during simulated icing.


NACA TN 2490
National Advisory Committee for Aeronautics.
FLIGHT INVESTIGATION OF SOME FACTORS AF-
FECTING THE CRITICAL TAIL LOADS ON LARGE
AIRPLANES. Harvey H. Brown. September 1951.
119p. diagrs., photos., 2 tabs. (NACA TN 2490)

Measurements were made of the control motions and
control forces during maneuvers regarded as likely
to produce critical tail loads. The pitching motion
of the airplane agreed well with that calculated, pro-
vided the equation of motion was not overly simpli-
fied. Structural deflections measured during the
maneuvers proved to be small. It was found that
pilots were inclined to allow the controls to return to
neutral at a higher rate than obtained in the initial
deflection, producing tail loads of which they had no
physical awareness.


NACA RM E51G02
National Advisory Committee for Aeronautics.
EXPERIMENTAL INVESTIGATION OF FORCED-
CONVECTION HEAT-TRANSFER CHARACTERIS-
TICS OF LEAD-BISMUTH EUTECTIC. Bernard
Lubarsky. September 1951. 30p. diagrs., photo.,
tab. (NACA RM E51G02)

The forced-convection heat-transfer characteristics
of lead-bismuth eutectic were experimentally inves-
tigated. Experimental values of Nusselt number for
lead-bismuth fell considerably below predicted
values. The addition of a wetting agent did not
change the heat transfer characteristics.



NACA RM E51G03
National Advisory Committee for Aeronautics.
ESTIMATION OF NEUTRON ENERGY FOR FIRST


RESONANCE FROM ABSORPTION CROSS SECTION
FOR THERMAL NEUTRONS. Donald Bogart.
September 1951. 18p. diagrs., 4 tabs. (NACA RM
E51G03)

Examination of published data for some 52 isotopes
indicates that the neutron energy for which the first
resonance occurs is related to the magnitude of the
thermal absorption cross section. The empirical
relation obtained is in qualitative agreement with the
results of a simplified version of the resonance
theory of the nucleus of Breit-Wigner.


NACA RM E51G05
National Advisory Committee for Aeronautics.
ADAPTATION OF A CASCADE IMPACTOR TO
FLIGHT MEASUREMENT OF DROPLET SIZE IN
CLOUDS. Joseph Levine and Kenneth S. Kleinknecht.
September 1951. 28p. diagrs., photos. (NACA RM
E51G05)

A cascade impactor, an instrument for obtaining the
size distribution of droplets borne in a low- velocity
air stream, has been adapted for flight cloud droplet
studies. Data from two flights are presented.


NACA TM 1274
National Advisory Committee for Aeronautics.
BEHAVIOR OF FAST MOVING FLOW OF COM-
PRESSIBLE GAS IN CYLINDRICAL PIPE IN
PRESENCE OF COOLING. (K Voprosu o
Povedenii Bystrodvizhushchegosya Potoka Szhimae-
mogo Gaza v Pryamoi Tsilindricheskoi Trube pri
Nalichii Okhlazhdenia). G. A. Varshavsky. Sep-
tember 1951. 8p. diagrs. (NACA TM 1274. Trans.
from Zhurnal Tekhnicheskoi Fiziki, v.16, no.4, 1946,
p.413-416).

For compressible flow with friction in a cylindrical
pipe the momentum, continuity, and heat-transfer
equations are examined to determine whether an in-
crease in Mach number ('thermal' Laval nozzle) is
obtainable through heat conduction from the gas
through the pipe walls. The analysis is based on the
assumption that the wall temperature is negligibly
small in comparison with the stagnation temperature
of the gas. The analysis leads to a negative
result. When the gas cooling is increased by also
considering radiation to the wall, a limited region
at high temperatures is obtained where Mach number
increases were theoretically possible. Obtaining
this condition practically is considered impossible.



BRITISH REPORTS


N-10001*

Aeronautical Research Council (Gt. Brit.)
FURTHER WIND TUNNEL TESTS ON THE EF-
FECTS OF ICE ACCRETION ON CONTROL CHAR-
ACTERISTICS. A. Spence. 1951. 12p. diagrs.,
2 tabs. (ARC CP 43; ARC 13,259. Formerly RAE
Tech. Note Aero 2048)

Describes tests on a tailplane with a 9-percent-thick
section, 12 trailing-edge angle and aspect ratio of 3





NACA
RESEARCH ABSTRACTS NO. 8


to find the effect of reducing the thickness-chord
ratio as well as the trailing-edge angle on the effects
of ice accretion on the control hinge moments. Re-
sults show that reduction of thickness has little ef-
fect but that the effects of transition movement on
b2(b2 = aCH/a,, where CH is the hinge-moment
coefficient and i is the flap angle) are of great
importance, especially at large trailing-edge angles.


N-10002*

Aeronautical Research Council (Gt. Brit.)
WIND TUNNEL TESTS ON THE EFFECT OF AC-
CRETION OF ICE ON CONTROL CHARACTERIS-
TICS IN TWO-DIMENSIONAL FLOW. A. S. Halliday,
A. S. Batson and D. K. Cox. (Revision of 12,403,
S June 9, 1949). May 1, 1950. 18p. diagrs., photo.
(ARC CP 42; ARC 13,112)

Tests in two-dimensional flow were made to provide
data for estimating the effect of ice accretion on a
tailplane when the trailing-edge angle of the airfoil
section is changed. The effects of aspect ratio on
the results are computed and plotted. Results show
that for elevators of 40% chord no advantage can be
expected from a finer trailing-edge angle; but for a
20% flap the finer trailing-edge appears to be supe-
rior. Larger trailing-edge angles, however, give
rise to much larger changes in control characteris-
tics due to normal shifts of transition point on the
tailplane and to changes in Reynolds number.


N-10003*

Aeronautical Research Council (Gt. Brit.)
A RECORDING SYSTEM FOR FLIGHT TEST DATA.
P. A. Hufton, F. G. R. Cook and P. S. Saunders.
1950. 13p. diagrs., photos. (ARC CP 44; ARC
12,925. Formerly AAEE Rept. A.A.E.E..'Res 246)

A new recording system for flight test data is pro-
posed in which the reading of the record is mecha-
nized. The system employs a transmission system
from the instruments which enables the instrument
readings to be recorded in the binary digital system.
This record is subsequently transcribed in an auto-
matic "reader" the output of which may be either in
the form of tabulated results, or as punched tape
suitable for direct use in a sequence-controlled cal-
culator. A prototype transmission unit has been
constructed.



N-10004*

Aeronautical Research Council (Gt. Brit.)
COMPARISON BETWEEN EXPERIMENTAL
MEASUREMENTS AND A SUGGESTED FORMULA
FOR THE VARIATION OF TURBULENT SKIN-
FRICTION IN COMPRESSIBLE FLOW. R. J.
Monaghan. 1951. 23p. diagrs. (ARC CP 45; ARC
13,260. Formerly RAE Tech. Note Aero 2037)

In ARC CP 46, a formula was suggested for the vari-
ation of turbulent skin friction with Mach number and
heat transfer. Experimental results gave a check on
the worth of the formula only for flow over a flat
plate under zero heat-transfer conditions at M = 2.46.
The present report extends this check by analyzing


data from Univ. of Tex., DRL CM-501, covering flow
over a flat plate under zero heat-transfer conditions
for M = 1.9 to M = 2.2, and from NACA RM E8L03,
covering subsonic flow through a circular pipe for
temperature differences up to 684 C. In each (over)
case, a good correlation is obtained on the basis of
a known incompressible flow formula.


N-10005*

Aeronautical Research Council (Gt. Brit.)
A NOTE ON THE USE OF TIME SERIES IN THE
ANALYSIS OF FLIGHT TEST RECORDS. W. P.
Jones. 1951. 28p. diagrs. (ARC CP 46; ARC
12,911)

Tustin's method for analyzing the behavior of linear
systems is briefly reviewed, the numerical analysis
being expressed in terms of matrices. Possible ap-
plications of time series representation to the study
of aircraft stability characteristics are discussed,
and a detailed numerical investigation of a simple
one degree of freedom undamped system is made.
For this system the Tustin method of analysis in
terms of A units seems satisfactory. An alterna-
tive method based on the use of Simpson's integra-
tion rule in conjunction with time series represen-
tation is also described.


N- 10006*

Aeronautical Research Council (Gt. Brit.)
THE EFFICIENCY OF AIRSCREWS ON WINGS OF
LARGE CHORD. E. Ower and R. Warden. 1951.
18p. diagrs., 4 tabs. (ARC R & M 2438. Formerly
ARC 3945; ARC Ae.Techl. 1509; ARC AP 191)

Presents information on the effect on propulsive ef-
ficiency of varying the distance between the plane of
the airscrew and the wing leading edge, when the
airscrew diameter is half the wing chord and the
wing thickness is sufficient to enclose the engine
nacelle. Gives the increase of lift due to the slip-
stream and compares results with calculations made
with the formula of Smelt and Davies. It also gives
a theoretical analysis of the effect on propulsive ef-
fiency of increasing the number of tractor airscrews
along the wing leading edge.



N-10007"

Aeronautical Research Council (Gt. Brit.)
CALCULATION OF THE INTERFERENCE ON A
THIN SYMMETRICAL AEROFOIL WITH HINGED
FLAP SPANNING A CLOSED WIND-TUNNEL. J. H.
Preston and A. R. Manwell. 1951. 14p. diagrs.
ARC R & M 2465. Formerly ARC 5388; ARC
Ae.1867)

Calculations of the tunnel interference corrections
for a2 aCLa"!, b1 = aCHi/a', b2 = aCH/an,
ml = aCm/aa', m2 = aCm/N (where Cm is the
moment coefficient about the leading edge, CH the
hinge moment coefficient, q the flap angle, and a the
incidence), were made for an infinitely thin symmet-
rical airfoil with hinged flap, spanning the center of
a closed wind tunnel. The method of solution is
analogous to that used by Glauert in R & M 1095, and





NACA
RESEARCH ABSTRACTS NO. 8


includes his results. The corrections for a2, b2,
and m2 are roughly equal to or less than that (over)
found by previous investigators for the lift, depending
on E the flap/chord ratio. The correction for ml
is about 50 percent greater and for b1 it is double.



N-10008*

Aeronautical Research Council (Gt. Brit.)
PITOT-TUBE READINGS NEAR SHOCK WAVES IN
THE N.P.L. AND R.A.E. HIGH SPEED TUNNELS.
R. G. Fowler. 1951. 13p. diagrs., photos. (ARC
R & M 2468. Formerly ARC 8383; ARC FM 773)

It is shown that in recent tests in the Royal Aircraft
Establishment High Speed Tunnel pitot readings near
shock waves agree with theory as well as would be
expected. Comparative readings in the National
Physical Laboratory 20" x 8" tunnel confirmed that
the condensation of atmospheric humidity has a large
but unpredictable effect on pitot pressures in this
tunnel and in all tunnels of this type.


N-10009*

Aeronautical Research Council (Gt. Brit.)
DISTORTION OF CONTROL SURFACE PANELS.
D. M. A. Leggett and R. G. Chapman. 1951. 30p.
diagrs., photos., 10 tabs. (ARC R & M 2478; ARC
8309; ARC 9991. Formerly RAE SME 3288; SME
3381)

Provides data for the design of control surface pan-
els of any required rigidity, flight experience having
shown that distortion of control surface panels, due
to air loads, often has a serious effect on the pilot's
operating force in the case of large or high-speed
aircraft. Part I contains estimates of the distortion
that is likely to occur with different types of con-
struction and for various aerodynamic loading. Part
II consists of a detailed account of tests made on
various types of control surfaces.


N-10010*

Aeronautical Research Council (Gt. Brit.)
WING FLEXURE-AILERON FLUTTER TESTS ON A
MODEL OF B.A.C. WING TYPE 167. C. Scruton.
1951. 19p. diagrs., photos., 8 tabs. (ARC R & M
2480; ARC 7566; ARC 8106)

Test results indicate that symmetrical wing flexure-
aileron flutter characteristics were slightly more fa-
vorable with a normal aileron control system than
with a proposed flexurally geared aileron system,
and the effect of mobility of the fuel in an unbaffled
tank was slightly beneficial. It was found also that
the minimum damping coefficient necessary to pre-
vent wing flexure-aileron flutter of the B.A.C. wing
type 167 varied from a value of 85 lb ft per radian
per second at sea level to 420 lb ft per second at an
altitude of 30,000 ft.


N-10011*

Aeronautical Research Council (Gt. Brit.)
NOTE ON DIFFERENTIAL GEARING AS A MEANS


OF AILERON BALANCE. S. B. Gates. 1951. 22p.
diagrs. (ARC R & M 2526; ARC 4921. Formerly
RAE BA 1642)

It has been suggested in some American investiga-
tions that differential gearing, combined with adjust-
ment of the aileron floating angle by means of a tab,
may be a powerful method of balancing ailerons.
This report sets out the theory of this method of bal-
ance and analyzes it in relation to the most pressing
problem of aileron design, which is to obtain close
balance at high speed without overbalance in any part
of the range, or uncomfortable lightness at slow
speed. It is shown that this result can be achieved
more directly by differential balance than by any
other method if the differential and the tab setting
are nicely adjusted to the natural floating properties
of the aileron.


N-10012*

Aeronautical Research Council (Gt. Brit.)
AN EXPERIMENTAL INVESTIGATION OF THE EF-
FECT OF LOCALISED MASSES ON THE FLUTTER
OF A MODEL WING. N. C. Lambourne and
D. Weston. 1951. 28p. diagrs., photos., 2 tabs.
(ARC R & M 2533. Formerly ARC 7604)

This report contains an account of some experiments
on the effect of concentrated masses (representing
wing engines, etc.) on the flutter characteristics of a
model cantilever wing. Flutter critical speeds and
frequencies were measured for an extensive range of
mass loading and the results are presented in the
form of diagrams. The flutter motions for a few
representative conditions of mass loading were de-
termined by an analysis of cinematograph pictures.
The results of experiments on the influence of the
flexibility of an engine mounting are also included.


N-10013*

Aeronautical Research Council (Gt. Brit.)
CALIBRATION OF THE ROYAL AIRCRAFT ESTAB-
LISHMENT 24-FT WIND TUNNEL. 1. E. Allen and
K. V. Diprose. 1951. 17p. diagrs., photo., 6 tabs.
(ARC R & M 2566; ARC 5980. Formerly RAE Tech.
Note 965)

A new calibration of the RAE 24-foot wind tunnel is
presented. The velocity distribution across the jet
in three planes was found at several tunnel speeds.
The distribution of static pressure throughout the
jet, the relation between the dynamic head at various
positions in the tunnel, and the hole-in-side pressure
were investigated. A complete list of all previous
calibrations together with results and reasons for
discrepancies is included. The plane of reference is
taken as 12.5 feet from the jet face, and the mean ve-
locity over the section is greater than the standard
value used up to the present by about 1/2 percent at
high speeds and 2 percent at low speeds.



N-10014*

Aeronautical Research Council (Gt. Brit.)
A COMPARISON BETWEEN PLAIN AND STRINGER-
REINFORCED SHEET FROM THE SHEAR LAG






NACA
RESEARCH ABSTRACTS NO. 8 7


STANDPOINT. M. Fine. 1951. 5p. diagrs. (ARC
R &M 2648. Formerly RAE SME 3182)

In R. &M's. 2098, 2099, 2100 the stringer-sheet meth-
od of solving shear lag problems in stringer rein-
forced sheet was developed. The present report
compares for two simple cases the solution for the
plain sheet with that for the stringer-reinforced
sheet. The solutions are practically identical by the
two methods provided that the sheet is considered
fully effective in taking end load. This leads to the
conclusion that, in regions of tensile stress, at all
events, all the skin area is to be included in the
stringer area when applying this method.



N-10120*

Royal Aircraft Establishment (Gt. Brit.)
A COMPARISON OF LEAD ACID AND ALKALINE
BATTERIES FOR USE IN M.T. VEHICLES. J. W. T.
Durrell and S. Pomroy. February 1951. 15p.
diagrs., tab. (RAE Tech. Note Chem.1144)

A comparison has been made of the capacity and
voltage characteristics of 24-volt M.T. batteries,
comprising standard lead acid and alkaline cells
ranging from 40 to 110 A.H. capacity at the 10-hour
rate, at temperatures of +20, 0, and -10 C and at
low and high discharge rates. The weight and vol-
ume of a lead acid battery are always considerably
less than those of an alkaline battery of equivalent
capacity and high rate voltage performance. The
comparison is unchanged even if the batteries are
half discharged before being subject to high rate
discharges.



UNPUBLISHED PAPERS



N-9423*

THE EFFECTIVE WIDTH OF INFINITELY LONG,
FLAT RECTANGULAR PLATES UNDER VARIOUS
CONDITIONS OF EDGE RESTRAINT. (De meedra-
gende breedte bij groote overschrijding der
knikspanning voor verschillende inklemming der
plaatranden). W. T. Koiter. 90p. diagrs., curves,
4 tabs. (Trans. from Nationaal Luchtvaart-
laboratorium, Amsterdam. S.287, December 31,
1943).

This investigation deals with the effective width of
(infinitely) long plates with straight, laterally freely
movable edges under axial compression. It is as-
sumed that the proportional limit is not exceeded and
that the plate does not buckle between the nails by
which it is attached to the stiffeners. The discussion
includes hinged, clamped, and elastically restrained
edges. In the last case, the torsional rigidity of the
stiffeners which produces the edge restraint is
taken as independent of the stress and of the wave
length of the buckles in the plate. The results ob-
tained for an infinitely long plate can also be applied
to plates of finite but not abnormally small length.


N-10135*

ON THE PROBLEM OF HYDROFOILS FOR SHIP
PROPULSION. (Kvoprosu o dvizhenii na podvodnykh
kryl'yakh). A. N. Vladimirov. 20p. diagrs., photos.
(Trans. from Sudostroenie, v.8, no.7, July 1938,
p.411-417).

In the journal Sudostroenie', No. 2, 1938, there ap-
peared an article by the engineering designer U. U.
Benua on certain problems on the propulsion of ships
by hydrofoils. The author gives some historical
data, points out the advantages that may be theoreti-
cally expected from the utilization of hydrofoils, and
presents his own proposal along ideas agreeing with
the proposal of the French engineer Grunberg. The
purpose of the present paper is to extend and deepen
somewhat the questions touched upon in the article
mentioned and to be of some assistance to those
working independently in this field in correctly ori-
enting themselves and soberly evaluating the actual
possibilities of hydrofoils.



N -10138"

BOUNDED FLOW WITH SEPARATION ABOUT A
CIRCULAR CYLINDER. (Otryvnoe Obtekanie
Kruglogo Tsilindra v Ogranichennom Potoke). Y. R.
Berman. lOp. diagrs. (Trans. from Prikladnaya
Matematika i Mekhanika, v.13, no.5, 1949,
p.543-546).

The two-dimensional problem of separated flow
about a symmetrical contour approximating a circle,
the flow being bounded by two parallel walls, is con-
sidered in this report. The flow scheme and the
conditions imposed are the usual ones for the prob-
lems of the classical flow theory.


N-10139*

NONLINEAR BOUNDARY PROBLEM OF UNSTEADY
MOTION OF VISCOUS INCOMPRESSIBLE FLUID.
(Nelineinaya Kraevaya Zadacha Neustanovivshegosya
Dvizhennia Vyazkoi Neszhimaemoi Zhidkosti). D. E.
Dolidze. 27p. (Trans. from Prikladnaya
Matematika i Mekhanika, v.12, no.2, Mar.-Apr.,
1948, p.165-180).

A viscous incompressible fluid filling an interior or
exterior region D bounded by the surface F is as-
sumed in a state of unsteady motion. The region D
is considered simply connected and unchanging in
time, and the surface F, regular; that is, surface F
is considered to have a continuously varying tangent
surface and principal curvature. This report con-
siders the problem of finding within the region D
regular solutions of the Navier-Stokes equations for
given values of the velocity on the boundary and at
the initial instant. By constructing the
so-called Green Tensor a nonlinear system of inte-
gral equations is obtained. The solution of this sys-
tem is effected by the method of successive approxi-
mations. A sufficient condition of convergence is
derived and a uniqueness theorem proved.


NACA-Langtey 10-1-51 2550




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