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
Publication Date:
Frequency:
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:
AA00009235:00001

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Sr.
K'


NO.30


CURRENT NACA REPORTS


NACA Rept. 1024
National Advisory Committee for Aeronautics.
CALCULATION OF THE LATERAL CONTROL OF
SWEPT AND UNSWEPT FLEXIBLE WINGS OF AR-
BITRARY STIFFNESS. Franklin W. Diederich.
1951. ii, 19p. diagrs., 6 tabs. (NACA Rept. 1024.


Formerly RM L8H24a) .

A method similar to that of NACA Rept. 1000 is -\
sented for calculating the effectiveness and the re \`
versal speed of lateral-control devices on swept a '
unswept wings of arbitrary stiffness. Provision is
made for using either stiffness curves and root-
rotation constants or structural influence coefficients
in the analysis. Computing forms and an illustrative
example are included to facilitate calculations by
means of the method. The effectiveness of conven-
tional aileron configurations and the margin against
aileron reversal are shown to be relatively low for
swept wings at all speeds and for all wing plan forms
at supersonic speeds.


NACA Rept. 1046
National Advisory Committee for Aeronautics.
A GENERAL INTEGRAL FORM OF THE BOUNDARY-
LAYER EQUATION FOR INCOMPRESSIBLE FLOW
WITH AN APPLICATION TO THE CALCULATION OF
THE SEPARATION POINT OF TURBULENT BOUND-
ARY LAYERS. Neal Tetervin and Chia Chiao Lin.
1951. 19p. diagrs. (NACA Rept. 1046. Formerly
TN 2158)

A general integral form of the boundary-layer equa-
tion, valid for either laminar or turbulent incom-
pressible boundary-layer flow, is derived. By using
the experimental finding that all velocity profiles of
the turbulent boundary layer form essentially a
single-parameter family, the general equation is
changed to an equation for the space rate of change of
the velocity-profile shape parameter. The lack of
precise knowledge concerning the surface shear and
the distribution of the shearing stress across turbu-
lent boundary layers prevented the attainment of a
reliable method for calculating the behavior of turbu-
lent boundary layers.


OCTOBER 9, 1952


It


'AVAILABLE ON LOAN ONLY.
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National Advisory Committee for Aeronautics


Research Abstracts


NACA Rept. 1056
National Advisory Committee for Aeronautics
THEORETICAL ANTISYMMETRIC SPAN LOADING
FOR WINGS OF ARBITRARY PLAN FORM AT SUB-
SONIC SPEEDS. John DeYoung. 1951. 36p.
diagrs., 9 tabs. (NACA Rept. 1056. Formerly
TNS 401

Ai hi ed lifting-surface theory that includes
effects of, ompressibility and spanwise variation of
secjsh lificurve slope is used to provide charts
with which antisymmetric loading due to arbitrary
antisy m-neric angle of attack can be found for wings
having'sya metric plan forms with a constant span-
wise'sweep angle of the quarter-chord line. Aero-
-- -dyniami 'characteristics due to rolling, deflected ai-
leroan, and sideslip of wings with dihedral are con-
-sidered. Solutions are presented for straight-tapered
wings for a range of swept plan forms.


NACA Rept. 1066
National Advisory Committee for Aeronautics.
ANALYSIS OF TEMPERATURE DISTRIBUTION IN
LIQUID-COOLED TURBINE BLADES. John N. B.
Livingood and W. Byron Brown. 1952. ii, 21p.
diagrs. (NACA Rept. 1066. Formerly TN 2321)

Analytical methods are presented for computing tem-
perature distributions in liquid-cooled turbine blades
or in simplified shapes used to approximate sections
of liquid-cooled turbine blades. Nondimensional
charts are presented for use in the simplification of
some of the calculations. Illustrative examples are
also included to demonstrate the use of the various
equations -and nondimensional charts and to show
trends of the various temperature distributions.
One-dimensional spanwise temperature distributions
gave satisfactory results near the coolant passages.
One-dimensional chordwise distribution gave a good
first approximation to the actual solution in cases
when rim cooling was insignificant.


NACA TN 2734
National Advisory Committee for Aeronautics.
SUMMARY OF AVAILABLE HAIL LITERATURE AND
THE EFFECT OF HAIL ON AIRCRAFT IN FLIGHT.
Robert K. Souter and Joseph B. Emerson.
September 1952. 162p. diagrs. ., photos., 6 Labs.
(NACA TN 2734)

Available information on the hail phenomenon affect-
ing aircraft in flight has been examined. This paper
attempts to coordinate the present knowledge of hail









2


with the effect of hail on aircraft in flight and in-
cludes (1) a digest of the literature on the physical
properties, the occurrence, and the formation of
hail; (2) a survey of the hail effect on aircraft in
flight from analyses of 57 cases of airplanes dam-
aged by hail, (3) a resume of hail information for the
benefit of pilots, forecasters, and ground operational
personnel; and (4) an annotated hail bibliography of
552 articles for use of research personnel



NACA TN 2765

A FLIGHT INVESTIGATION OF THE EFFECT OF
SHAPE AND THICKNESS OF THE BOUNDARY LAY-
ER ON THE PRESSURE DISTRIBUTION IN THE
PRESENCE OF SHOCK. Eziaslav N. Harrin.
September 1952. 13p. diagrs., photos. (NACA
TN 2765)

An investigation was made in flight at free-stream
Mach numbers up to about 0. 77 to determine the ef-
fect of a laminar boundary layer and thin and thick
turbulent boundary layers on the chordwise pressure
distribution over an airfoil in the presence of shock
at full-scale Reynolds numbers. Boundary-layer
and pressure-distribution measurements were made
on a short-span airfoi built around a wing of a
fighter airplane. Boundary-layer Reynolds numbers
(based on momentum thickness and flow parameters
at the outer edge of the boundary layer) were about
3,000 for the laminar boundary layer and 10, 000 for
the thickest turbulent boundary layer with local Mach
numbers ranging up to 1. 3 and chord Reynolds num-
bers up to about 21 x 106.


NACA TN 2773

AN APPROXIMATE METHOD FOR DETERMINING
THE DISPLACEMENT EFFECTS AND VISCOUS
DRAG OF LAMINAR BOUNDARY LAYERS IN TWO-
DIMENSIONAL HYPERSONIC FLOW. Mitchel H.
Bertram. September 1952. 41p. diagrs., photos.,
tab. (NACA TN 2773)

A simplified approximate theory is presented by
means of which the laminar boundary layer over an
insulated two-dimensional surface may be calculated,
a linear velocity profile being assumed, and an esti-
mate made of its effect in changing the pressure
distribution over the profile upon which the boundary
layer is formed. Skin friction is also determined.
Comparisons of results from this theory are made
with experimental results at a Mach number of 6. 86
and a Reynolds number of 980,000.


NACA TN 2780

FLIGHT INVESTIGATION OF TRANSIENT WING RE-
SPONSE ON A FOUR-ENGINE BOMBER AIRPLANE
IN ROUGH AIR WITH RESPECT TO CENTER-OF-
GRAVITY ACCELERATIONS. Harry C. Mickleboro,
Richard B. Fahrer and C. C. Shufflebarger.
September 1952 25p diagrs., 3 taos. (NACA
TN 2780)


NACA
RESEARCH ABSTRACTS Nd

The results of a flight investigation on a four-engine
bomber airplane to determine the transient-response
effects of wing flexibility in gusts showed that the
measured acceleration increments at the center of
gravity were approximately 28 percent higher than
the true airplane acceleration increments. This re-
latlonship appeared to be only slightly affected by the
variations in speed and weight covered by the test
conditions.


NACA TN 2783

USE OF A CONSOLIDATED POROUS MEDIUM FOR
MEASUREMENT OF FLOW RATE AND VISCOSITY
OF GASES AT ELEVATED PRESSURES AND TEM-
PERATURES. Martin B. Biles and J. A. Putnam,
University of California. September 1952. 51p.
diagrs., photos., 7 tabs. (NACA TN 2783)

Use of a consolidated porous medium as a gas-
metering device and for determination of gas viscos-
ity has been investigated over a moderate range of
temperature and pressure. With normal laboratory
techniques it appears possible to calibrate large po-
rous Alundum filtering thimbles to meter gas with a
probable error of 0. 1 to 0. 2 percent. The geometry
of such elements permits an appreciable range of gas
flow rate to be metered with small, accurately con-
trolled. pressure drops. The advantages of such a
device warrant its use as a laboratory instrument.
Results of the flow tests have been employed in the
determination of the viscosity of air up to approdi-
mately 900 pounds per square inch absolute at the
two test temperatures of 750 and 5170 F. These
data appear to check sufficiently well with other
published viscosity data to justify the use of this
method as a recommended procedure.

NACA TN 2784

METHOD FOR CALCULATION OF COMPRESSIBLE
LAMINAR BOUNDARY-LAYER CHARACTERISTICS
IN AXIAL PRESSURE GRADIENT WITH ZERO HEAT
TRANSFER. Morris Morduchow and Joseph H.
Clarke, Polytechnic Institute of Brooklyn.
September 1952. 43p. diagrs., 4 tabs. (NACA
TN 2784)

The Karmin-Pohlhausen method Is extended primari-
ly to sixth-degree velocity profiles for determining
the characteristics of the compressible laminar
boundary layer over an adiabatic wall in the presence
of an axial pressure gradient. It is assumed that the
Prandtl number is unity and that the coefficient of
viscosity varies linearly with the temperature. A
general approximate solution which permits a rapid
determination of the boundary-layer characteristics
for any given free-stream Mach number and given
velocity distribution at the outer edge of the boundary
layer is obtained. A simple method based on the use
of a seventh-degree velocity profile is derived for the
special purpose of calculating the location of the sep-
aration point in an adverse pressure gradient. It is
shown that for the special case of flow near a forward
stagnation point the Karmdn-Pohlhausen method with
the usual forth-degree profiles leads to results of
adequate accuracy, even for the critical Reynolds
number.









NACA
RESEARCH ABSTRACTS N0.30


NACA TN 2785

INTRODUCTION TO ELECTRICAL-CIRCUIT ANAL-
OGIES FOR BEAM ANALYSIS Stanley U. Benscoter
and Richard H. MacNeal, California Institute of
Technology. September 1952. 48p. diagrs., 5 tabs.
(NACA TN 2785)


An application is described of the well-known analogy
between electrical and mechanical systems to the
calculation of stresses and deflections of beams.
The object of the present paper is to give an explana-
tion of the analogies in an elementary manner which
will enable a structural engineer to understand the
process of designing the electrical circuits. The
analogies which are discussed are those that are now
being used in the Cal-Tech analog computer. Anal-
ogies are given for beams in bending and torsion with
static loads and in vibrational motion.



NACA TN 2786

EQUIVALENT PLATE THEORY FOR A STRAIGHT
MULTICELL WING. Stanley U. Benscoter and
Richard H. MacNeal, California Institute of Technol-
ogy. September 1952. 32p. diagrs. (NACA
TN 2786)

A structural theory is developed for the analysis of
thin multicell wings with straight spars and perpen-
dicular ribs. The analysis is intended to be suitable
for supersonic wings of low aspect ratio. Deflections
due to shearing strains are taken into account. The
theory is expressed entirely in terms of first-order
difference equations in order that analogous electrical
circuits can be readily designed and solutions obtained
on the Cal -Tech analog computer.


NACA TN 2787

AIRFOIL PROFILES FOR MINIMUM PRESSURE
DRAG AT SUPERSONIC VELOCITIES APPLICA-
TION OF SHOCK-EXPANSION THEORY, INCLUD-
ING CONSIDERATION OF HYPERSONIC RANGE.
Dean R. Chapman. September 1952. 44p. diagrs.
(NACA TN 2787)

A theoretical investigation employing shock expan-
sion theory is made of the airfoil profile having
minimum pressure drag at zero lift for various given
auxiliary conditions. Curves are presented to facili-
tate the application of the theory, and typical opti-
mum profiles are illustrated. A comparison of re-
sults obtained by the simpler linearized-theory meth-
od of a previous report indicates that the simpler
method can be used with engineering accuracy to de-
termine the shape, though not the drag, of the opti-
mum profile at Mach numbers up to infinity. It is
shown that considerable deviation from the optimum
shape can be made without a large increase in drag
except on thin airfolls at moderate supersonic Mach
numbers.


3


NACA TN 2788

EFFECTS OF SOLVENTS IN IMPROVING BOUNDARY
LUBRICATION OF STEEL BY SILICONES. S. F.
Murray and Robert L. Johnson. September 1952
23p. diagrs., 2 tabs. (NACA TN 2788)

Because of the known synthetic fluids, silicones best
satisfy the viscometric requirements for lubricants
for turbine engines, a study was conducted to estab-
lish the effect of solvents on boundary lubrication by
silicones. Boundary-lubrication data were obtained
which are considered substantiating evidence for a
hypothesis that, in solutions of solvents blended with
silicones, the silicones form a closely packed and
oriented adsorbed film on ferrous surfaces. The so-
lutions reduced friction and prevented surface'failure
even when the solvent as well as the silicone was an
extremely poor lubricant. These data indicate that
satisfactory lubrication is the result of a solvation
effect rather than a lubrication additive effect of the
solvent because 30 to 50 percent of solvent was nec-
essary for good results. The best results were ob-
tained with solvents having dipole moments. Solu-
tions of diesters in silicones may be practical'jubri-
cants.


NACA TN 2789

SOME DYNAMIC EFFECTS OF FUEL MOTION IN
SIMPLIFIED MODEL TIP TANKS ON SUDDENLY
EXCITED BENDING OSCILLATIONS. Kenneth F.
Merten and Bertrand H. Stephenson. Septemtler
1952. 35p. diagrs., photos., 2 tabs. (NACA
TN 2789)

Arnexploratory investigation of the dynamic effects of
fuel sloshing in tip tanks on wing bending motion was
conducted with two simplified model beam-tank sys-
tems. Envelope curves to beam -displacement time
histories obtained after release from a deflected
position are compared and show the effects of yaria-
tion in tank fullness, fluid density, fluid viscosity,
and tank shape. Some variations of fluid weight ef-
fective from cycle to cycle are also presented? The
results of these tests indicate that after several
cycles substantial damping may be obtained fearn
fuel sloshing in a tip tank and that the effective mass
of the fuel may vary considerable under certaj4 con-
ditions of tank oscillation. The viscosity of the fluid
did not affect the damping or inertia characteristics
obtained but, for a given beam-tank system, the
density of fluid and tank fullness were important
parameters.



NACA TN 2791

CORRELATION OF TENSILE STRENGTH, TENSILE
DUCTILITY, AND NOTCH TENSILE STRENGTH
WITH THE STRENGTH OF ROTATING DISKS OF
SEVERAL DESIGNS IN THE RANGE OF LOW AND
INTERMEDIATE DUCTILITY Arthur G. Holms and
Andrew J. Repko. September 1952. 30p diagrs.,
3 tabs (NACA TN 2791)







4



Burst tests were conducted on several designs of
sound disks and disks with defects and results were
compared with tensile strength, tensile ductility, and
notch tensile strength. For the brittle materials, the
disk strength did not correlate with tensile strength.
For the brittle materials and for ductile materials
for which notch strength data were available, the disk
strength was found to correlate better with the com-
bination of tensile strength and notch strength ratio
than with the combination of tensile strength and elon-
gation. For disks with defects, the notch tensile
strength was superior to the conventional tensile
strength.


NACA TN 2793

A METHOD FOR THE DETERMINATION OF THE
TIME LAG IN PRESSURE MEASURING SYSTEMS
INCORPORATING CAPILLARIES. Archibald R.
Sinclair and A. Warner Robins. September 1952.
35p. diagrs., tab. (NACA TN 2793)

A method is presented for the determination of the
time lag in pressure measuring systems incorporat-
ing capillaries; this method is a convenient and sys-
tematic means of selecting, designing, or redesign-
ing a pressure measuring system to meet the time
requirements of a particular Installation. Experi-
mental data are shown and a step-by-step sample
application is presented.


NACA TN 2795

EFFECTS OF WING SWEEP ON THE UPWASH AT
THE PROPELLER PLANES OF MULTIENGINE AIR-
PLANES. Vernon L. Rogallo. September 1952.
46p. diagrs. (NACA TN 2795)

An analysis is presented to give a qualitative picture
of the effects of wing sweep on the upwash at the
propeller planes of multiengine airplanes. In order
to provide a basis for judging effects of sweep, com-
parisons are made of the upwash and upflow angles at
the propeller planes of two hypothetical airplanes of
the high-speed long-range type, one having an un-
swept wing and the other a sweptback wing. The
effects of compressibility are considered Charts
are provided to enable the prediction of upwash in the
chord-plane region ahead of wings of various plan
forms.



NACA RM E52B12

IMPINGEMENT OF WATER DROPLETS ON AN
NACA 651-212 AIRFOIL AT AN ANGLE OF ATTACK
OF 40. Rinaldo J Brun, John S Serafini and
George J. Moshos. September 1952. 47p. diagrs.,
tab. (NACARM E52B12)

The trajectories of droplets in the air flowing past an
NACA 651-212 airfoil at an angle of attack of 40 were
determined. The collection efficiency, the area of
droplet impingement. and the rate of droplet impinge-
ment were calculated from the trajectories. The re-
sults are applicable under the following conditions:


NACA
RESEARCH ABSTRACTS NO.30



chord lengths from 2 to 20 feet, altitudes from 1000
to 35. 000 feet, airplane speeds from 150 miles per
hour to the critical night Mach number, and droplet
diameters from 5 to 100 microns.



NACA RM E52H15

PRESSURE LIMITS OF FLAME PROPAGATION OF
PURE HYDROCARBON-AIR MIXTURES AT RE -
DUCED PRESSURES. Adolph E. Spakowski.
September 1952. 35p. diagrs., 2 tabs. (NACA
RM E52H15)

An investigation was made of the pressure concen-
tration limits of flame propagation in glass tubes for
18 high-boiling hydrocarbons mixed with air. The
concentration limits were correlated with molecular
weight. Relations were also derived between the
flammability range and the molecular weight and be-
tween the rich and lean limits.


NACA RM L52G18

INVESTIGATION OF THE EFFECTS OF VARIATIONS
IN THE REYNpLDS NUMBER BETWEEN 0. 4 X 106
AND 3. 0 X 10 ON THE LOW-SPEED AERODY-
NAMIC CHARACTERISTICS OF THREE LOW-
ASPECT-RATIO SYMMETRICAL WINGS WITH
RECTANGULAR PLAN FORMS. George W. Jones,
Jr. September 1952. 13p. diagrs. (NACA
RM L52G18)

The effects of Reynolds number on the aerodynamic
characteristics of three symmetrical wings of aspect
ratio 1. 2, and 3 each having a rectangular plan
form and an NACA 0012 airfoil section is given for a
range of sevgn Reynolds numbers between 0. 4 x 106
and 3. 0 x 10 The data show the effects of
Reynolds number on the lift, the lift-curve slope,
maximum lift coefficient, and the pitching moment
about the quarter-chord point.


NACA TM 1337

ANALYTICAL STUDY OF SHIMMY OF AIRPLANE
WHEELS. (Etude Th6orlque du Shimmy des Roues
d'Avlon). Christian Bourcier de Carbon. September
1952 126p. diagrs., photos. (NACA TM 1337.
Trans. from Office National d'Etudes et de
Recherches Aeronautiques, Pub. 7, 1948).

The problem of shimmy of a casterini, wheel, such as
the nose wheel of a tricycle gear airplane, is treated
analytically. The flexibility of the tire is considered
to be the primary cause of shimmy. The rather
simple theory developed agrees rather well with pre-
vious experimental results. The author suggests
that shimmy may be eliminated through a suitable
choice of landing gear dimensions in lieu of a damper.






NACA
RESEARCH ABSTRACTS NO.30



BRITISH REPORTS


N-17089,

Aeronautical Research Council (Gt. Brit.)
TWO-DIMENSIONAL WIND TUNNEL INTERFER-
ENCE. L. G Whitehead. June 17, 1950. 24p.
diagrs. (ARC 13, 198; FM 1451)

The present paper describes a contribution to the
theory of two-dimensional wind-tunnel interference
on bodies set at zero incidence in the center of the
flow. The work is limited to incompressible flow.
The irrotational flow has been found between paral-
lel walls and also in a free jet past two cylindrical
sections. The first of these is a circular cylinder
and the second is a slender profile with a uniform
pressure drop over the greater part of the boundary.
No account is taken of the wake which exists behind
a body in a real fluid so that the tunnel interference
represented in the present paper is that usually re-
ferred to as solid blockage and the effects of wake
blockage are not considered. The free streamline
method is employed for the calculation of the flow of
a jet past both bodies so that the outward displace-
ment of the streamline is correctly represented and
the approximations associated with the application of
the method of images to these problems are avoided.
The results are not limited therefore to examples in
which the profile dimensions are small compared
with the width of the jet. The free streamline meth-
od has an additional advantage as the solutions can
readily be adapted to deal with problems in which the
outer boundaries of the flow consist partly of paral-
lel walls and partly of free streamlines.

N-17091*

Aeronautical Research Council (Gt. Brit.)
THE LAMINAR AX]-SYMMETRIC JET: EXACT
SOLUTION. H. B. Squire. July 22, 1950. 7p.
diagrs. (ARC 13, 267; FM 1462)

A solution of the flow in a lamminar axisymmetric jet
was given by Schlichting in 1933, making use of the
approximations of boundary-layer theory. The cor-
responding exact solution of the Navler-Stokes equa-
tions is derived in the present report. It is also
shown that this solution can be interpreted as the
flow resulting from a force applied at a point in a
viscous fluida which Is at rest at infinity.


N-17099'

Forest Products Research Lab. (Gt. Brit.)
TRIALS OF TIMBER FOR PLYWOOD MANUFAC-
TURE. PRELIMINARY REPORT ON SIX AFRICAN
SPECIES PROGRESS REPORT SEVENTEEN.
June 1952 28p. 15 tabs. IForest Products Re-
search Lab.)

This report describes tests on six African species of
wood to determine their suitability for plywood man-
ufacture. The following species were tested: ber-
linia, brown sterculia, kokrodua, mguma, pterygota,
and yellow sterculia.


5


N-17109*

National Gas Turbine Establishment (Gt. Brit.)
SOME EXPERIMENTS ON BREAKDOWN POTENTIAL
IN HOT PRODUCTS OF COMBUSTION. N. A.
Dimmock. May 1952. 14p. diagrs., ab. (NGTE
Memo. M. 153)

The experiments described in this memorandum show
that no general lowering of the breakdown potential is'
caused by the ionization present in the exhaust gases
from a combustion chamber, although a slight drop in
Its value is noticeable very close to the chamber exit.
The fuels used were aviation kerosene, gas oil
(Marine reference B. 310) and gas oil with n-Butyl
disulphide. The ceramic insulator used in the ex-
periments was found to become semiconducting at
elevated temperatures and this trouble would have to
be overcome in designing an electrode for a hot gas
electrostatic -cleaner.


N-17172'

Royal Aircraft Establishment (Gt. Brit.)
MEASUREMENT OF THE MOISTURE CONTENT OF
HIGH PRESSURE OXYGEN FOR USE IN AIRCRAFT.
I. A LOW PRESSURE HYGROMETER. M E
Bedwell and W. G. Glendinning. March 1952. 9p.
diagrs., 3 tabs. (RAE Tech. Note Chem. 11541

A hygrometer has been devised which will give pre-
cise measurement of the moisture content of oxygen
issuing from a cylinder. Jest runs on samples from
transport cylinders indicate that in every case the
oxygen contained considerably less than the maximum
amount of moisture allowed by specification. No
anomalous changes of moisture content with decreas-
ing pressure were observed in the cylinders selected.


N-17304*

Royal Aircraft Establishment (Gt. Brit.)
FREQUENCY REGULATOR FORA MOTOR-
ALTERNATOR C. S. Hudson. January 1952.
17p. diagrs., photo. (RAE Tech. Note EL. 30)

The design of a frequency regular for a 2 kw 333 cps
motor alternator is described and details given of
the experimental measurements made. A two stage
magnetic amplifier is used having a current gain of
about 3000 andia power amplification of about 108.
The design is intended to give regulation to within
0. 1 percent; from approximate measurements made
it is considered to be better than 0.3 percent. No
attempt has been made to examine the stability of the
frequency reference circuit and further work on this
needs to be done.


N-17305'

Royal Aircraft Establishment (GI Brit.)
SOME MEASUREMENTS OF TAKEOVER TIME LAGS
IN THE STRIKING OF COLD-CATHODE TRIODES.
J. C. LeGrlce. March 1952. 16p. diagrs., photos.
(RAE Tech. Note EL. 32)




4* 4


6


The design of time-delay circuits using cold-cathode
triodes necessitated a knowledge of the magnitude of
the takeover time by which striking of the anode cir-
cuit lags behind striking of the trigger circuit. A
method of applying the trigger and anode voltage
wave forms to the deflection plates of a C. R. 0. is de-
scribed which enabled the takeover time to be mea-
sured between the attainment of currents of 10-3
amps in the trigger and anode circuits respectively.
Some measurements on miniature triodes with potas-
sium cathodes gave takeover times of less than 50
Ms for anode voltages between 100 and 200 volts.
The takeover times were repeatable to better than
5 percent.


N-17314*

Royal Aircraft Establishment (Gt. Brit.)
TOWING TANK TESTS TO DETERMINE THE WATER
DRAG AND PITCHING MOMENTS ON THE FINAL
HULL FORM OF A LARGE FLYING BOAT SEA-
PLANE (PRINCESS, SPEC. 10,'46). T. B. Owen
and A. G. Kurn. April 1952. 33p. diagrs., photo.,
5 tabs. (RAE Tech. Note Aero 2159)

The water drag, trim and hull wetted areas of the
Princess flying boat hull model (final form) have
been measured for all anticipated take-off weights
and altitudes. The drag characteristics compare
favorably with those obtained on other boat seaplanes.
The measurements were made using new techniques
for: model surface finish, drag measurements, and
representation of the effect of air flow past the model.
A method is given for correcting the model drag data
to full scale to allow for the difference is Reynolds
number.


NACA .
RESEARCH ABSTRACTS NO.30, j


UNIVERSITY OF FLORIDA


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NACA-Langley 10-9-52




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