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:
AA00009235:00035

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National Advisory Committee for Aeronautics


Research Abstracts


NO.72

CURRENT NACA REPORTS
NACA Rept. 1150

CONSIDERATIONS ON THE EFFECT OF WIND-
TUNNEL WALLS ON OSCILLATING AIR FORCES
FOR TWO-DIMENSIONAL SUBSONIC COMPRES-
SIBLE FLOW. Harry L. Runyan and Charles E.
Watkins. 1953. ii, 7p. diagrs. (NACA Rept. 1150.
Formerly TN 2552)

This paper treats the effect of wind-tunnel walls on
the oscillating two-dimensional air forces in a com-
pressible medium. The walls are simulated by the
usual method of placing images at appropriate dis-
tances above and below the wing. An important re-
sult shown is that, for certain conditions of wing fre-
quency, tunnel height, and Mach number, the tunnel
and wing may form a resonant system so that the
forces on the wing are greatly changed from the con-
dition of no tunnel walls.


NACA Rept. 1162

LIFT DEVELOPED ON UNRESTRAINED RECTAN-
GULAR WINGS ENTERING GUSTS AT SUBSONIC
AND SUPERSONIC SPEEDS. Harvard Lomax. 1954.
ii, 16p. diagrs., 5 Labs. (NACA Rept. 1162.
Formerly TN 2925)

Lift forces induced by a vertical gust are estimated
on the basis of theoretical calculations. The effects
of pitching and wing bending are neglected and only
wings of rectangular plan form are considered.
However, the effects of Mach number (from 0 to 2)
and aspect ratio (2 to ) are included, and solutions
are given by means of which the response to gusts
having arbitrary streamwise gradients can be cal-
culated. Results are presented for sharp-edged
and triangular gusts and various wing-air density
ratios.



NACA TM 1376

EXPERIMENTS ON TAIL-WHEEL SHIMMY.
(Experimentelle Untersuchungen uber das
Spornradflattern). 0. Dietz and R. Harling.
October 1954. 81p. diagrs., photos. (NACA
TM 1376. Trans. from Zentrale fiur
weissenschaftliches Berichtswesen der
Luftfahrtforschung, Berlin, FB 1320)

Model tests on the "running belt' anJ tests with a
full-scale tail wheel were made on a rotating drum as
well as on a runway in order to investigate the causes
of the undesirable shimmy phenomena frequently oc-
curring on airplane tail wheels, and the means of


worn.OCTOBER 27, 1954 |

avoiding them. The smallmodel (sraem1:Th) per-
matted simulation of the mass, moments of inertia,
and fuselage stiffnesses of the airplane and determi-
nation of their influence on the shimmy, whereas by
means of the larger model with pneumatic tires
(scale 1:2) more accurate investigations were made
on the tail wheel itself. The results of drum and
road tests show good agreement with one another and
with the model values. Detailed investigations were
made regarding the dependence of the shimmy
tendency on trail, rolling speed, load, size of tires,
ground friction, and inclination of the swivel axis;
furthermore, regarding the influence of devices with
restoring effect on the tail wheel, and the friction
damping required for prevention of shimmy. Finally
observations from slow-motion pictures are reported
and conclusions drawn concerning the influence of
Lire deformation.



.

NACA TN 3245

CALCULATED SUBSONIC SPAN LAJASANT ./<
RESULTING STABILITY DERIVAT V&FN
SWEPT AND 450 SWEPTBACK TAIL SURFACES IN
SIDESLIP AND IN STEADY ROLL. M. J. Queijo
and Donald R. Riley. October 1954. HO0p.
diagrs., 2 tabs. (NACA TN 3245)
Subsonic span loads have been calculated for a sys-
tematic series of unswept and 450 sweptback tail
surfaces by the discrete-horseshoe-vortex method.
Results are presented as charts of span loads and
the resulting contribution to the side-force and
rolling derivatives. Geometric variables covered in
the investigation included vertical- and horizontal-
tail aspect ratios and vertical position and dihedralof
the horizontal tail for various tail assemblies inside-
slip and steady roll. Also presented is an extensive
table of values of sidewash due to a rectangular
vortex.





NACA TN 3247

AN EVALUATION OF AN ACCELEROMETER
METHOD FOR OBTAINING LANDING-GEAR DRAG
LOADS. Jerome G. Theisen and Philip M. Edge. Jr.
October 1954. 22p. diagrs., photos. (NACA
TN 3247)

An evaluation is made of applied ground drag loads
obtained by means of angular-acceleration measure-
ments on the wheel of a landing gear together with


*AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST, NW., WASHINGTON s25. D. C.. CITING CODE NUMBER ABOVE EACH TITLE.
THE REPORT TITLE AND AUTHOR.

(aZF. /30of

2/r-5







2

vertical-acceleration measurements on the upper and
lower masses. The data are obtained during forward-
speed landing impacts and drop tests in the Langley
impact basin. The results obtained by using this
method showed good agreement with values obtained
simultaneously from specially constructed dynamom-
eters.



NACA TN 3248

AN EXPERIMENTAL AND THEORETICAL INVESTI-
GATION OF THE ANISOTROPY OF 3S ALUMINUM-
ALLOY SHEET IN THE PLASTIC RANGE. Arthur J.
McEvily, Jr. and Philip J. Hughes. October 1954.
45p. diagrs., photos., 4 tabs. (NACA TN 3248)

The results of tension and compression tests on 3S
aluminum-alloy sheet are presented together with
values of Poisson's ratio in the plastic range. Crys-
tallographic anisotropy was found to be responsible
for the variation in Poisson's ratio in the width and
thickness directions of specimens alined at various
directions with respect to the rolling direction in the
plane of the sheet. X-ray diffraction studies of
annealed specimens were made and pole figures
showing the type of preferred orientation were drawn.
A theoretical analysis was made to account for the
observed anisotropy based on the behavior of single
crystals and the texture of the sheet. An independent
check of the theory is afforded through comparison
with experimental data on cubically alined copper
sheet.


NACA TN 3249

THE HYDRODYNAMIC CHARACTERISTICS OF AN
ASPECT-RATIO-0.125 MODIFIED RECTANGULAR
FLAT PLATE OPERATING NEAR A FREE WATER
SURFACE. John A. Ramsen and Victor L. Vaughan,
Jr. October 1954. 32p. diagrs. (NACA TN 3249)

Results of an investigation of the hydrodynamic force
characteristics of an aspect-ratio-0.125 modified flat
plate on a single strut operating near a free water
surface are presented. Comparisons between these
data and similar data from previous tests on plates
having aspect ratios of 1.00 and 0.25 are presented;
these comparisons give the effects of aspect ratio on -
the characteristics. The experimental effects of
varying the depth of submersion and angle of attack
are shown. The effects of cavitation at the leading
edge and of the planing-bubble type of high-angle
separation are presented. Comparisons are also pre-
sented between the experimental lift for all three
aspect ratios at large depths ot submersion and
several theoretical methods of predicting the lift.


NACA TN 3250

AN EXPERIMENTAL INVESTIGATION OF THE
EFFECT OF WHEEL PREROTATION ON LANDING-
GEAR DRAG LOADS. Dexter M. Potter. October
1954. 19p. diagrs., photos., tab. (NACA TN 3250)

The effect of prerotation on the wheel spin-up drag
loads on a small landing gear during landings has
been investigated in the Langley impact basin. The
results showed that, at the forward speed of the test
(approximately 85 fps), even partial prerotation
resulted in a reduction in both the drag load and the
vertical load. At very high speeds, however, con-


NACA
RESEARCH ABSTRACTS NO.72
sideration of the variation of drag load with forward
speed as obtained in a previous investigation indicat-
ed that large amounts of prerotation must be used in
order to assure a reduction in drag load.


NACA TN 3263

LIFT AND MOMENT EQUATIONS FOR OSCILLAT-
ING AIRFOILS IN AN INFINITE UNSTAGGERED
CASCADE. Alexander Mendelson and Robert W.
Carroll. October 1954. 46p. diagrs., 3 tabs.
(NACA TN 3263)

Aerodynamic coefficients similar to those of the
isolated airfoil are obtained as functions of the cas-
cade geometry and the phasing between successive
blades; the phasings considered are zero, 900, and
1800. These aerodynamic coefficients are plotted
for the special case when all the airfoils are vibrat-
ing in bending in phase (3600 phasing). It is shown
that the effect of cascading for this case is to reduce
greatly the aerodynamic damping.



NACA TN 3265

VAPORIZATION RATES AND DRAG COEFFICIENTS
FOR ISOOCTANE SPRAYS IN TURBULENT AIR
STREAMS. Robert D. Ingebo. October 1954. 39p.
diagrs., photos., 3 tabs. (NACA TN 3265)

Drop-size distribution and drop-velocity data were
obtained for isooctane sprays in turbulent air streams
using a droplet camera developed at the NACA Lewis
laboratory. Experimental spray vaporization rates,
based on the mean diameter, correlated single-
droplet vaporization rates. An empirical expression
was derived for isooctane droplet drag coefficients.




BRITISH REPORTS




N-32700*

Ministry of Supply (Gt. Brit.)
THE APPLICATION OF "REDUX" ADHESIVE TO
AIRCRAFT STRUCTURES. April 1954. 157p.
diagrs., photos., 37 tabs. (MOS S & TM 5/54)

This report is concerned with the use of Redux as
an adhesive in typical aircraft structures. The
strength of the Redux joint was tested by means
of static and fatigue tests after considering the
variables of pretreatment, pressure, and tempera-
ture conditions.



N-32702*

Ministry of Supply (Gt. Brit.)
REPORT ON THE TUNGSTEN ARGON-ARC WELD-
ING OF MAGNESIUM-ZIRCONIUM ALLOYS. AN
INVESTIGATION INTO THE USE OF MAGNESIUM
ZIRCONIUM ALLOYS IN SHEET AND PLATE
FORM. April 1954. 103p. diagrs., photos., 21 tabs.
(MOS S & TM 6/54)






NACA
RESEARCH ABSTRACTS NO. 72
This report is an investigation as to the suitability
of magnesium-zirconium alloys fabricated by argon
arc welding for aircraft structures. The strength
of these sheet alloy welds was tested by means of
compression and tension loads.




N-32751

Aeronautical Research Council (Gt.Brit.)
AN EXPERIMENTAL STUDY OF THREE-
DIMENSIONAL HIGH-SPEED AIR CONDITIONS IN A
CASCADE OF AXIAL-FLOW COMPRESSOR BLADES.
K. W. Todd. 1954. 34p. diagrs., photos. (ARC
R & M 2792; ARC 12,711; 12,308)

Detailed investigations have been made by optical
and physical methods in a high-speed wind tunnel of
the flow characteristics of two compressor cas-
cades. In part 1 a representative high-camber cas-
cade was examined at zero incidence over entry air
velocities ranging from low to critical. Traverses
were made of discharge angles and wake losses at
all heights so that a relation between two- and three-
dimensional losses could be obtained. In part 2 the
blade and passage design was conditioned by the find-
ings of part 1, with the aim of so modifying the cas-
cade that its efficiency in the critical flow region
would be improved.


N-32752

Aeronautical Research Council (Gt. Brit.)
A NEW LAW OF SIMILARITY FOR PROFILES,
VALID IN THE TRANSONIC REGION.
K. Oswatitsch. 1954. 12p. diagrs. (ARC
R & M 2715; ARC 10,807. Formerly RAE Tech.
Note Aero 1902)

A new law of similarity is given, valid for slender
profiles in mixed transonic flow with negligible
viscosity, according to which the cube of the
Prandtl factor of any critical Mach number is
proportional to the thickness ratio. It is shown
that this rule, and that of von Kirman for flow at
sonic speed, are valid for shock waves within the
range over which the shock loss is proportional to
the cube of the pressure rise. Experimental
pressure distributions plotted according to this
rule show good agreement, except for the position
of the shock wave on the surface.



N-32754 ,

Aeronautical Research Council (Gt.Brit.)
TREATMENT OF THE STAGNATION POINT IN
ARITHMETICAL METHODS. A. Thorn. 1954. 0lp.
diagrs., 2 tabs. (ARC R &M 2807. Formerly
ARC 14,118; FM 1577; Oxford Univ., Engineering
Lab. No.53)

In using any of the relaxation techniques near a
stagnation point difficulties arise if the variable is
log 1. q. This tends to infinity and the difference
equation no longer represents adequately the
differential equation without special modification.
Methods are given whereby larger squares can be
used than had been previously practicable. As little


variation as possible has been introduced into the
procedure so that if an electronic calculator is
used, the alterations to the circuits would be reduced
to a minimum.


N-32755*

Aeronautical Research Council (Gt.Brit.)
VELOCITY DISTRIBUTION ON THIN TAPERED
WINGS WITH FORE-AND-AFT SYMMETRY AND
SPANWISE CONSTANT THICKNESS RATIO AT
ZERO INCIDENCE. S. Neumark and J. Colling-
bourne. 1954. 43p. diagrs., 6 tabs. (ARC
R & M 2858; ARC 14,474. Formerly RAE Aero
2432)

This report is a continuation of three earlier ones
by the present authors and contains a theoretical
investigation of subsonic flow past thin tapered
unswept wings (of full- or cropped-rhombus
plan form), at zero angle of attack. The method
used is that of linear perturbation in which the wing
is replaced by a system of sources and sinks; the
required velocity field is then approximately ob-
tained as that of the velocity components, in the
plane of the chord lines, given by this source
system. The supervelocity Is then determined by
direct integration.



N-32756 *

Aeronautical Research Council (Gt.Brit.)
METHODS OF APPROACHING AN ACCURATE
THREE-DIMENSIONAL POTENTIAL SOLUTION
FOR A WING. H. C. Garner. 1954. 19p. diagrs.,
2 tabs. (ARC R & M 2721. Formerly ARC 11,802;
Perf.489; FM 1293)

Attention is given to some fundamental aspects of
the vortex-sheet theory for determining the distri-
bution of lift on a finite wing. The accuracy and
limitations of some existing approximate forms of
the theory are discussed. With special reference
to the labor of computation an iterative approach
to an accurate solution is suggested, and the general
mathematical expression for the distribution of
lift required to give an exact solution for a Vee wing
is considered.



N-32830*

Royal Aircraft Establishment (Gt.Brit.)
TRANSDUCTORS WITH VOLTAGE RE-SET
CONTROL. A. G. Milnes. March 1954. 21p.
diagrs. (RAE Tech.Note EL.67)

Voltage reset transductors have half-cycle delay
times but this response is at the expense of very
limited amplification. Factors influencing the per-
formance are discussed and a voltage feedback
technique is examined. When preamplification is
required thermionic triodes, transistors or con-
ventional transductors may be used; but the control
voltage and current conditions of a reset circuit
may present matching problems. Applications as
computing elements, as synchro-coupling units, as
lead networks and for multiplication of two voltages
are discussed.










N-32832 *

Royal Aircraft Establishment (Gt.Brit.)
HEAT TRANSFER FROM CIRCULAR FINNED
CYLINDERS UNDER NATURAL CONVECTION
CONDITIONS. Mary D. Wright and G. A. French.
March 1954. 17p. diagrs., 6 tabs. (RAE Tech.
Note EL.68)

An experimental investigation has recently been
made into the cooling in free air of cylinders
fitted with different sizes and thicknesses of
circular fins and with varied fin spacings.




N-32836*

Royal Aircraft Establishment (Gt.Brit.)
DESIGN OF TRIDAC (A THREE-DIMENSIONAL
SIMULATOR FOR GUIDED MISSILES). PART VI -
MONITORING TECHNIQUES. K. C. Garner. May
1954. 18p. diagrs., photos. (RAE Tech.Note
GW 318)

Monitoring systems have been developed to detect
certain fault conditions in the operation of the
majority of the individual components comprising
TRIDAC. Voltage variations or failures in the
power supplies to the d.c. amplifiers, and abnormal
drift voltages which may occur in the d.c. amplifiers
are suitably indicated. Monitoring is also provided
to detect saturation of the signal output from each
of the d.c. amplifiers. The location in the simulator
at which any of these faults occur is displayed on a
system of warning lights. The basic principles of
these, and other associated facilities, considered of
importance to fault location and operational proce-
dure, are discussed.




N-32860*

Aeronautical Research Council (Gt. Brit.)
EXPERIMENTS ON THE FLOW INTO A SWEPT
LEADING-EDGE INTAKE AT ZERO FORWARD
SPEED WITH NOTES ON THE WIDER USES OF A
SLOTTED INTAKE. J. Seddon and W. J. G.
Trebble. 1954. 18p. diagrs., tab. (ARC
R & M 2909; ARC 13, 928. Formerly RAE
Aero 2409)

The flow into a swept intake at zero forward speed
(ground running conditions) is shown to be analogous
to the flow round a sharp corner in a duct. Tests
have been made on a model of a swept-wing leading-
edge intake to measure the losses involved. It is
found that the distribution inside the duct can be
improved by the use of straight guide vanes,
alternatively by means of a special intake slot, or
further by a combination of both. Guide vanes
increase the mean loss, but the intake slot improves
(i. e., reduces) this also. The slot would require to
be sealed under flight conditions. It is suggested
that this form of slotted intake may have wider
applications in the future. Using the results of the
experiments and an analogy with the slotted wing,
conclusions are drawn regarding the main points
of design of the intake slot.


NACA
RESEARCH ABSTRACTS NO. 72

N-32862*

Aeronautical Research Council (Gt. Brit.)
THE ROLLING POWER OF AN ELASTIC SWEPT
WING. E. G. Broadbent. 1954. 19p. diagrs.,
4 tabs. (ARC R &M 2857; ARC 13,639. Formerly
RAE Structures 85)

An iterative method of solution is given for the
problem of loss in rolling power due to wing defor-
mation. The method is applicable at subsonic or
supersonic speeds, and compressibility effects are
allowed for, provided the variation of the aerody-
namic derivatives with Mach number is known. The
numerical labor involved in the solution is not
great and the accuracy is considerably greater than
can be achieved by the semirigid method





N-32866*

Nat. Gas Turbine Establishment (Gt.Brit.)
A THEORETICAL EXPRESSION FOR POINT-
SOURCE DIFFUSION IN TURBULENT FLOW. A. B.
P. Beeton. March 1954. 18p. diagrs. (NGTE
R. 152)

A general expression governing the turbulent dif-
fusion of fuel from a point source in a uniform
stream is established from theoretical considera-
tions. The distribution at points some distance
downstream of the source can then be reduced to a
relatively simple equation, which is identical in form
with the relation found to hold experimentally for
any one section normal to the flow. It indicates a
slight modification to the normal empirical law
relating distributions at different sections, again
confirming the results of experimental work. The
general expression enables distributions to be
predicted for points as close to the source as may
be desired. In comparison with the accurate values
obtained by graphical integration of this expression,
the approximate equation is found to give better than
1 percent accuracy at all points beyond about 5 mix-
ing lengths downstream of the source.





N-32868*

Royal Aircraft Establishment (Gt. Brit.)
FREQUENCY TRIPLING OF AN ELECTRIC SUPPLY
BY SATURABLE TRANSFORMER AND TRANS-
DUCTOR TECHNIQUES. F. D. Gill. March 1954.
52p. diagrs., photos., 2 tabs. (RAE Tech. Note
EL. 62)

Known methods of frequency tripling by means of
saturable transformers are examined and found to be
too sensitive to input voltage variations for some
purposes. A tranaductor technique has been de-
veloped which will tolerate substantial input voltage
variations (8 percent). The method is sensitive to
the value of load impedance, which must be kept
reasonably constant. The frequency tripler is found
to be suitable as a power supply to a magnetic ampli-
fier. Both the weight and response time of the mag-
netic amplifier, redesigned for the triple frequency,






NACA
RESEARCH ABSTRACTS N0.72

are reduced. In a typical example it is shown that
the reduction in amplifier weight compensates for
the weight of the frequency tripler and the additional
power requirements. The principal advantage of the
technique is the improvement in amplifier response
tune by a factor of three.



N-32869[

Royal Aircraft Establishment (Gt.Brit.)
A CUP-TYPE TRANSMITTING ANEMOMETER
WITH IMPROVED RESPONSE TO DECREASING
GUSTS. W. J. G. Cox. July 1954. 28p. diagrs.,
photo. (RAE Tech.Note Instn. 140)

A cup-type anemometer is described, capable of
giving remote indication and recording of air speed.
The instrument is self-generating, giving a voltage
output which is reasonably proportional to air speed.
A damping brake is described which automatically
comes into operation only while the speed is decay-
ing, thereby minimizing the inherent free wheel
tendency of this type of instrument and improving
its follow-up characteristic to winds of decreasing
velocity. The transient responses of the braked
and free rotor to step function air velocity changes
are deduced, and calibration and loading curves for
a particular model are given.



N-32870

Royal Aircraft Establishment (Gt.Brit.)
A RANGE OF 50 C/S TRANSDUCTORS. A. G.
Milnes and W. N. Corn. April 1954. 23p. diagrs.,
3 tabs. (RAE Tech.Note EL.64)

Design details are given for twelve 50 c/s trans-
ductors covering a power range up to 200 VA at the
limit of linear response. The cores used are of
Mumetal or H.C.R. laminations or Hipersil C-cores
and the voltage range is 13, 25, 50, 115, 240 and
420 V (rms). Nondimensional transductor charac-
teristics are obtained from the performance curves
of these units and applied in a design example.



N-32871'

Royal Aircraft Establishment (Gt.Brit.)
DEVELOPMENT OF A VERY QUICK ACTING
ELECTRO HYDRAULIC VALVE. F. W. Read.
June 1954. 12p. diagrs., photo. (RAE Tech.Note
LAP 1030)

A quick acting hydraulic "on-off" valve was re-
quired by Aero Department, R.A.E. as part of ex-
perimental equipment to limit the stresses in the
tail structure of an aircraft fitted with a powered
elevator control and a "g" restrictor. In the first
place a "Dowty" type valve was modified and an
"on-off, cycle was completed in 45 milliseconds at
700 psi. Flight trials, however, indicated that a
shorter operating time was necessary and this note
describes a new design in which full hydraulic
pressure operates the valves to both "close" and
"open" position. A notable improvement in operat-
ing times, in rig tests at 450 psi, was shown by the
valve "closing" or "opening" in 6 to 7 milliseconds.


N-32876*

Aeroplane and Armament Experimental Establish-
ment (Gt. Brit.) AN ANALYSIS OF THE LONGITU-
DINAL STABILITY AND CONTROL OF A SINGLE
ROTOR HELICOPTER. F. O'Hara. July 16, 1954.
25p. diagrs. (AAEE/Res/280)

The general theory of longitudinal stability and con-
trol for a single rotor helicopter is presented in a
form similar to that for fixed wing aircraft. It is
shown to be possible to establish for the helicopter
in forward flight, in the same way as for fixed wing
aircraft, stick fixed static and maneuver margins,
on which the stability and handling qualities depend
to a marked extent. Extension of the theory is
required for stick free longitudinal stability and also
for rotor wing combinations and tandem rotor
helicopters.



N-32877*

Royal Aircraft Establishment (Gt.Brit.)
INTENSITY/TIME CHARACTERISTICS OF THREE
EMITTED WAVEBANDS IN A CAPACITOR DIS-
CHARGE THROUGH XENON. R. L. Aspden. May
1954. lOp. diagrs., photos. (RAE Tech.Note
Instn.139)

Intensity time oscillograms, recording the lumi-
nous emission obtained when a capacitor is dis-
charged through gas or vapor, conceal any time
factor which may be related to specific frequencies
or wavebands. The preliminary investigation de-
scribed herein shows that in the case of a low
pressure xenon filled tube, flashed at normal load-
ing, both the time to peak intensity and the pulse
duration are functions of frequency. Although the
light is a source of continuously varying spectral
quality the composite nature of the flash suggests
that for a very limited period during discharge the
emission may approach blackbody radiation. A
method whereby this period may be located and
used to obtain the optimum reproduction in color of
a high speed transient is outlined. Details are
given of the equipment used in the present tests,
together with intensity/time curves of three re-
corded wavebands covering the visible spectrum.





N-32889*

Aeronautical Research Council (Gt.Brit.)
A VARIABLE INDUCTANCE ACCELERATION
TRANSDUCER. H. K. P. Neubert. 1954. 19p.
diagrs., photos., 2 tabs. (ARC CP 169)

The acceleration transducer described has been
designed as a general purpose transducer for the
measurement of vibration and the resonance testing
of aircraft on the ground and in flight. The two
ranges of t3g and 9g, with cutoff frequencies at
70 c/sec and 100 c/sec respectively, will cover the
most common magnitudes of vibration amplitudes
and frequencies in these applications. It is designed
for use with a carrier bridge amplifier and galvanom-
eter recorder, providing a frequency range of zero
to about 100 c/sec.









N-32890*

Aeronautical Research Council (Gt.Brit.)
THE RAPID, ACCURATE PREDICTION OF
PRESSURE ON NON-LIFTING OGIVAL HEADS OF
ARBITRARY SHAPE, AT SUPERSONIC SPEEDS.
B. W. Bolton-Shaw and H. K. Zienkiewicz, English
Electric Co., Luton. 1954. 58p. diagrs., 3 tabs.
(ARC CP 154)

Five methods are developed for determining the
pressure distribution on an arbitrary, pointed, con-
vex axisymmetric head shape, of which the ordi-
nate and its first three derivatives are everywhere
continuous. When the geometric details are speci-
fied, the time required to pressure plot a head shape
is from 20 minutes to 3 hours. The methods were
checked in five trial cases against accurate pressure
distribution, obtained by using van Dyke's second
order theory, and generally gave very good agree-
ment. The best method to use of the five depends
upon head shape and on the speed and accuracy
required.



N-32891 *

Aeronautical Research Council (Gt.Brit.)
THE TECHNIQUE OF FLUTTER CALCULATIONS.
H. Templeton. 1954. 65p. diagrs., 22 tabs.
(ARC CP 172)

This report describes the basic principles on which
theoretical flutter analyses are made, and illus-
trates them by some simple applications. The
techniques employed are typical of those in current
use in this country. Three appendices give the two-
dimensional aerodynamic derivatives for a wing-
aileron-tab system, computational details of typical
forms of solution, and an illustration of the use of
resonance test modes in flutter calculations.



N-32893*

Aeronautical Research Council (Gt.Brit.)
NOTE ON AN APPLICATION OF THE TILTING
PLATE METHOD OF MACH NUMBER VARIATION
FOR WIND TUNNEL TESTS AT LOW SUPERSONIC
SPEEDS. J. Seddon and L. Haverty. 1954. 18p.
diagrs., photo. (ARC CP 168)

This report describes an application of the method
to a study of internal flow problems of side intakes
at transonic speeds in a small supersonic tunnel.
By arrangements involving the use of three or four
tunnel nozzles, a continuous Mach number range
from 0.5 to 1.6 is made available, apart from a gap
between 0.97 and 1.04. This method is being used
in the No. 2, 5-1/2 inch x 5-1/2 inch supersonic
tunnel of the R.A.E.



N-32894 *

Aeronautical Research Council (Gt.Brit.)
THE MANUFACTURE OF AEROFOIL MODELS BY
TANGENT PLANE MILLING. R. S. Marriner.
1954. 20p. diagrs. (ARC CP 166)


NACA
RESEARCH ABSTRACTS No.72

The method described in this paper is a milling
process followed by a minimum of hand finishing to
ensure a smooth blending of the tangential planes.
The accuracy achieved in this trial manufacture of
a half-delta wing model was generally just within
tO.001 inch. The geometrical ideas and formulas
developed are for a particular type of airfoil model,
but the general idea of rotating and tilting a blank
under a cutter can be applied to other types.




N-32895*

Aeronautical Research Council (Gt.Brit.)
SCHLIEREN METHODS FOR OBSERVING HIGH-
SPEED FLOWS. D. W. Holder and R. J. North.
1954. 27p. diagrs., photos. (ARC CP 167)

This paper discusses the present state of knowledge
concerning the use of schlieren and direct-shadow
methods for the visualization of high-speed flow, the
emphasis being on the use of the methods in wind
tunnel experiments. The techniques for observing
flows which are two-dimensional or which possess
axial symmetry have reached an advanced stage of
development, and are satisfactory for most investi-
gations of this kind; recent progress is reviewed.
Several methods which may be useful for the study
of flow round finite wings and wing-body combina-
tions are described briefly.




N-32896*

Aeronautical Research Council (Gt.Brit.)
THE THERMODYNAMICS OF FRICTIONAL RE-
SISTED ADIABATIC FLOW OF GASES THROUGH
DUCTS OF CONSTANT AND VARYING CROSS
SECTION. W. R. Thomson. 1954. 44p. diagrs.
(ARC CP 158)

This report presents an analytical study dealing with
the adiabatic flow of gases with frictional losses
through ducts of constant and varying cross section.
The thermodynamic treatment is along lines pub-
lished by other workers such as Bailey and Fabri
and is essentially one dimensional in character
insofar that frictional effects are assumed to be
uniformly distributed over the total cross sectional
area of flow. With this simplifying assumption,
relationships are deduced connecting the pressure,
temperature, velocity, and flow area of the gas at
any one plane with those at any other plane in a duct.




N-32897*

Aeronautical Research Council (Gt.Brit.)
UNSTEADY CAVITATING FLOW PAST CURVED
OBSTACLES. L. C. Woods. 1954. 9p. diagrs.
(ARC CP 149)

The plane incompressible flow past two symmetric
curved obstacles, between which is a finite constant
pressure cavity, is calculated for the case when
the cavity length and pressure are functions of time.







NACA
RESEARCH ABSTRACTS NO. 72

N-32898 *

Aeronautical Research Council (Gt. Brit.)
AN INVESTIGATION INTO THE ROLLING POWER
AND AILERON REVERSAL CHARACTERISTICS OF
SWEPT WINGS. A. V. Coles and R. J. Margetts.
1954. 28p. diagrs., photo., 3 tabs. (ARC CP 159)

This paper reviews briefly existing theories relating
to swept wings and presents the results obtained from
a program of tests on flexible swept wings. The
wings tested were of constant span, aspect ratio and
area and had angles of sweep ranging from 150 sweep-
forward to 600 sweepback. The tests were carried
out in a low-speed wind-tunnel, compressibility
effects being absent. The test results showed that
the reversal speeds of wings of the same stiffness
increase with both sweepback and sweepforward, the
minimum value occurring with a small amount of
sweepback. Predictions of reversal speeds based on
semirigid theory showed good agreement with
experimental results.




N-32899*

Aeronautical Research Council (Gt.Brit.)
THE PERFORMANCE OF A MULTI-ENGINE HELI-
COPTER FOLLOWING FAILURE OF ONE ENGINE
DURING TAKE-OFF OR LANDING. A. L. Oliver.
1954. 15p. diagrs., tab. (ARC CP 175)

A theoretical analysis has been made considering the
type of site proposed for civil operation in built-up
areas. The performance of a twin engine helicopter
similar to the Bristol 173 appears to be adequate for
safe operation, but the handling judgment involved
in return landings may make the performance diffi-
cult to achieve. A take-off technique allowing climb-
away after engine failure at any stage is preferable
but this is not possible for the twin engine machine
within the space available. It is possible if the twin
engines are replaced by four of the same effective
total power but only if the turning climb-away is
made after engine failure. A helicopter with suf-
ficient performance for a straight climb-away can
in general hover with one engine inoperative.




N-32900 *

Aeronautical Research Council (Gt.Brit.)
AERODYNAMIC DERIVATIVES FOR A DELTA WING
OSCILLATING IN ELASTIC MODES. D. L. Wood-
cock. 1954. 33p. diagrs., 13 tabs. (ARC CP 170)

Aerodynamic derivatives are given for a delta wing
of aspect ratio 3 and 900 apex angle oscillating with
symmetric elastic modes in incompressible inviscid
flow. They have been determined by the lattice
method of W. P. Jones, using the values of the down-
wash calculated by D. E. Lehrian when obtaining
aerodynamic derivatives for the same delta wing
oscillating in rigid wing modes. Results for several


7


modes are given both as local derivatives and also
as equivalent constant derivatives. Derivatives for
other modes can be obtained either from these or
from the values of the reciprocal of the downwash
matrix, which also are tabulated.



N-33258*

Aeronautical Research Council (Gt. Brit.)
TOWING-TANK TESTS ON A LARGE SIX-ENGINE
FLYING BOAT SEAPLANE, TO SPECIFICATION
10/46 (PRINCESS). PART H PORPOISING
STABILITY, SPRAY AND AIR DRAG TESTS, WITH
IMPROVED STEP FAIRING, AFTERBODY DESIGN
AND AERODYNAMIC MODIFICATIONS. A. G.
Smith, D. F. Wright and T. B. Owen. 1954. 32p.
diagrs., photos., 2 tabs. (ARC R & M 2834; ARC
13,964. Formerly RAE Aero 2404)
Tests were made to improve the main-step fairing
in order to reduce air drag while retaining satis-
factory porpoisingstabilityat high water speeds;
to reduce the midplaning porpoising instability found
with the hull lines tested in Part I; to test the effect
of increased wing and tailplane areas; and to predict
more accurately the full-scale performance of the
final hull form by representing more closely the
anticipated full-scale conditions of lift, slipstream
and damping in pitch. The final form evolved is
used for the first production aircraft.






MISCELLANEOUS





N-33233

Advisory Group for Aeronautical Research and
Development. THE MECHANISM OF CARBON
FORMATION. George Porter. (Scheveningen
Netherlands Conference May 3-7, 1954). 16p.
diagrs. (AG13/M9)

This paper gives a more detailed account of the
theory of carbon formation in combustion processes
which has been briefly stated elsewhere. Other
theories are critically reviewed and a summary is
given of the experimental data relating to carbon
formation which has been obtained by the flash
photolysis technique. The mechanism proposed,
which is believed to apply to the high temperature
pyrolysis of hydrocarbons as well as to diffusion
flames, premixed flames, and explosions in closed
vessels, is as follows: all hydrocarbons, when
pyrolysed under the high temperature conditions
prevalent in the above systems, undergo a
decomposition which results eventually in the for-
mation of acetylene and hydrogen. Acetylene then
forms carbon particles by a simultaneous conden-
sation and dehydrogenation.




UNIVERSITY OF FLORIDA


3 1262 08153 091 6


UNPUBLISHED PAPERS

N-23711*

SOME RESEARCH WORK IN PROGRESS ON TURBO-
MACHINES. (Quelques Recherches en Cours sur
Turbomachines). E. Maillet. May 1954. 26p.
diagrs., photos. (Trans. from Association Tech-
nique Maritime et Aeronautique, 1953)

Experimental studies carried out on a centrifugal
compressor and on components of axial-flow
compressors are discussed. Test methods, tech-
nique, and measuring equipment are described.




N-23775*

ON EXACT SOLUTIONS OF THE STOKES-NAVIER
EQUATIONS OF INCOMPRESSIBLE FLUIDS AT
MODIFIED BOUNDARY CONDITIONS. (Uber exakte
Losungen der Stokes-Navier-Gleichungen
inkompressibler Flussigkeiten bei verainderten
Grenzbedingungen). Jakob Ackeret. September 1954.
20p. diagrs. (Trans. from Zeitschrift fuir
angewandte Mathematik und Physik, v. 3, no.4, 1952,
p. 259-271)

There exist solutions of the hydrodynamic equations
if the boundary conditions are modified. If it can be
arranged that the "solid" boundaries are actually
moving with the fluid velocity then really simple
solutions are obtained. Several examples are worked
out. As turbulence is unlikely to occur under these
conditions, the power to move bodies in a viscous


NACA
RESEARCH ABSTRACTS NO. 72


fluid at high Reynolds numbers would become ex-
tremely low. It seems that even a partial fulfilment
of these changed conditions would be of practical
interest.






DECLASSIFIED NACA REPORTS








NACA RM L9B10

AN ANALYSIS OF AVAILABLE DATA ON EFFECTS
OF WING-FUSELAGE-TAIL AND WING-NACELLE
INTERFERENCE ON THE DISTRIBUTION OF THE
AIR LOAD AMONG COMPONENTS OF AIRPLANES.
Bertram C. Wollner. April 11, 1949. 33p. diagrs.,
tab. (NACA RM LB110) (Declassified from
Confidential, 8/18/54)

Available information on the effects of wing-fuselage-
tall and wing-nacelle interference on the distribution
of the air load among components of airplanes is
analyzed. The effects of several variables are con-
sidered. Of these, vertical wing position is found to
affect the change in lift caused by interference be-
tween wing and fuselage to the greatest extent.


NACA-Langley 10-26-54 4




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