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:
AA00009235:00013

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National Advisory Committee for Aeronautics.


Research Abstracts
NOVEMBER 17T 1953


NO. 52


CURRENT NACA REPORTS

NACA Rept. 1094

AN EXPERIMENTAL INVESTIGATION OF TRAN-
SONIC FLOW PAST TWO-DIMENSIONAL WEDGE
AND CIRCULAR-ARC SECTIONS USING A MACH-
ZEHNDER INTERFEROMETER. Arthur Earl
Bryson, Jr., California Institute of Technology.
1952. ii, 33p. diagrs., photos. (NACA Rept. 1094.
Formerly TN 2560)

Interferometer measurements are given of the flow
fields near two-dimensional wedge and circular-arc
sections at zero angle of attack. Pressure distribu-
tions and drag coefficients as functions of Mach num-
ber were obtained and the wedge data are compared
with theory. It is shown that the local Mach number
at any point on the surface of a finite three dimen-
sional body or an unswept two-dimensional body,
moving through an infinite fluid, has a stationary
value at Mach number 1 and, in fact, remains nearly
constant for a range of speeds below and above Mach
number 1. On the basis of this concept and the
experimental data, pressure distributions and drag
coefficients for the wedge and circular-arc sections
are presented throughout the entire transonic range
of velocities.


NACA Rept. 1097

STRESSES IN A TWO-BAY NONCIRCULAR CYLIN-
DER UNDER TRANSVERSE LOADS. George E.
Griffith. 1952. ii, 12p. diagrs.. 3 tabs. (NACA
Rept. 1097. Formerly TN 2512)

A method, taking into account the effects of flexibility
and based on a general eighth-order differential
equation, is presented for finding the stresses in a
two-bay, noncircular cylinder the cross section of
which can be composed of circular arcs. Numerical
examples are given for two cases of ring flexibility
for a cylinder of doubly symmetrical (essentially el-
liptic) cross section, subjected to concentrated
radial, moment, and tangential loads. The results
parallel those already obtained for shells with circu-
lar rings.

NACA Rept. 1112

HYDROCARBON AND NONHYDROCARBON DERIV-
ATIVES OF CYCLOPROPANE. Vernon A. Slabey,
Paul H. Wise and Louis C. Gibbons. 1953. ii,
18p. diagram 4 tabs. (NACA Rept. 1112)

The methods used to prepare and purify 19 hydro-
carbon derivatives of cyclopropane are discussed.
Of these hydrocarbons, 13 were synthesized for the
first time. In addition to the hydrocarbons, six
cyclopropylcarbinols, five alkyl cyclopropyl ketones,


three cyclopropyl chlorides, and one cyclopro-
panedicarboxylate were prepared as synthesis inter-
mediates. The melting points, boiling points, re-
fractive indices, densities, and, in some instances,
heats of combustion of both the hydrocarbon and
nonhydrocarbon derivatives of cyclopropane were
determined. These data and the infrared spectrum
of each of the 34 cyclopropane compounds are pre-
sented herein. The infrared absorption bands
characteristic of the cyclopropyl ring are discussed,
and some observations are made on the contribution
of the cyclopropyl ring to the molecular refractions
of cyclopropane compounds.


NACA Rept. 1113

sriCTRUM OF TURBULENCE IN A CONTRACTING
STREAM. H. S. Ribner and M. Tucker. 1953. ii,
17p. diagrs., tab. (NACA Rept. 1113. Formerly
TN 2606)

Spectrum concepts are employed to study the selec-
tive effect of a stream contraction on longitudinal and
lateral turbulent velocity fluctuations. By consider-
ation of the effect of stream contraction on a single
plane wave, the effect on spectrum and correlation
tensors of the turbulence is determined. Weak turbu-
lence and an inviscid fluid are postulated; compress-
ibility of mean flow only is taken into account. For
axisymmetric contraction and isotropic initial turbu-
lence, explicit results s'_ e-ebaLned. The one dimen-
sional longitudinal sf'trum 'ffitd to be markedly
distorted. The oa'eqhive efll re -i t reaction on lon-
gitudinal and lapra components of turbulence is found
to be given unique regardless a of the iso-
tropic spectrum; pomparisn 3 !peri ent ismade.

NACA TN 3022

METHOD FOR STUDYING HEyw6P'ER LONGI-
TUDINAL MANEUVER STATITLITY.' Kenneth B.
Amer. October 1953. f, 52p. *diagrs., photos.,
2 tabs. (NACA TN 3022)

A theoretical analysis of helicopter maneuver stabil-
ity is made and the results are compared with ex-
perimental results for both a single- and a tandem-
rotor helicopter. Techniques are described for
measuring in flight the significant stability deriva-
tives for use with the theory to aid in design studies
of means for achieving marginal maneuver stability
for a prototype helicopter.

NACA TN 3026

ELECTROSTATIC SPARK IGNITION-SOURCE
HAZARD IN AIRPLANE CRASHES. Arthur M. Busch.
October 1953. 28p. diagrs., photos., tab. (NACA
TN 3026)

The hazard of igniting airplane crash fires by elec-


* AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASHINGTON 2s, D.C., CITING CODE NUMBER ABOVE EACH TITLE,
THE REPORT TITLE AND AUTHOR.
eg.9, /Jo0?






2

trostatic sparks, generated when detached airplane
parts fly through clouds of dust and fuel mist, was
investigated. Within the limits of variables studied,
the rates with which airplane wreckage collected a
charge were directly proportional to the rate that
clay dust or fuel mist was intercepted. Maximum
rates of experimental electrification were used to
relate energy accumulation to wreckage sizes and
trajectories and to estimate minimum hazardous
wreckage sizes and trajectories. Comparison of
sizes and trajectories of wreckage shown in motion
pictures of airplane crashes with these estimated
sizes and trajectories indicated that the hazard is
small. Of the remedial measures considered, poly-
ethylene coatings were found to offer promise of
protection against electrostatic spark ignition.


NACA TN 3027

INFLUENCE OF ROTOR-ENGINE TORSIONAL
OSCILLATION ON CONTROL OF GAS-TURBINE
ENGINE GEARED TO HELICOPTER ROTORS. John
C. Sanders. October 1953. 40p. diagrs., photos.,
2 tabs. (NACA TN 3027)

Equations were developed for the torsional motion of
a gas-turbine engine geared to a helicopter rotor in
which the rotor blades were hinged to the rotor shaft.
The rotor system was simplified to yield simple
third-order equations that can be used in the analysis
of engine control. Comparison of the system re-
sponse calculated from these equations with the ex-
perimentally observed frequency response of a rotor
from a 2500-pound helicopter showed satisfactory
agreement. Calculations showed that the torsional
motion arising from the hinged construction of the
blades contributed to instability of the engine-speed
and engine-torque controls. Trends in stability and
response of controls with increasing weight of heli-
copters were investigated.


NACA TN 3028

THE COMPRESSIBLE LAMINAR BOUNDARY
LAYER WITH HEAT TRANSFER AND SMALL
PRESSURE GRADIENT. George M. Low.
APPENDIX B. NUMERICAL SOLUTION OF DIFFER-
ENTIAL EQUATIONS. Lynn U. Albers. October
1953. 68p. diagrs., 7 tabs. (NACA TN 3028)

A perturbation method for the calculation of velocity
and temperature profiles, skin-friction and heat-
transfer characteristics for two-dimensional com-
pressible laminar boundary layers with heat transfer
and small arbitrary gradient is presented. The
permissible pressure gradients include those of a
form and magnitude usually encountered over slender
aerodynamic shapes in supersonic flight. The
combined effects of heat transfer and pressure gradi-
ent on boundary-layer characteristics are demon-
strated by applying the results of the analysis to two
representative wings.

NACA TN 3032

AN ANALYTICAL STUDY OF THE EFFECT OF AIR-
PLANE WAKE ON THE LATERAL DISPERSION OF
AERIAL SPRAYS. Wilmer H. Reed, III. October
1953. 46p. diagrs., 3 tabs. (NACA TN 3032)

An analysis is made to determine the trajectories and
deposit of aerial spray droplets which are issued into
the air disturbances generated by an agricultural


NACA
RESEARCH ABSTRACTS NO. 52
airplane. Various nozzle arrangements and droplet-
size spectra are considered with a view to improving
the uniformity and effective width of the deposUt.
-4

NACA TN 3034

GRAPHICAL SOLUTION OF SOME AUTOMATIC-
CONTROL PROBLEMS INVOLVING SATURATION
EFFECTS WITH APPLICATION TO YAW DAMPERS
FOR AIRCRAFT. William H. Phillips. October
1953. 41p. diagrs. (NACA TN 3034)

A graphical method is presented for determining the
motion of a freely oscillating system of one degree of
freedom stabilized by a controlling device which
applies control force in proportion to the displace-
ment of the system, to its rate of change of dis-
placement, or both. The controlling member is
assumed to have limitations on its maximum deflec -
tion and on its maximum rate of movement. Several
examples concerned with the application of yaw
dampers to aircraft are presented to illustrate the
method, and some conclusions regarding the varia-
tion of the stability of the motion with amplitude are
obtained.

NACA TN 3035

A PRELIMINARY STUDY OF THE PROBLEM OF
DESIGNING HIGH-SPEED AIRPLANES WITH SATIS-
FACTORY INHERENT DAMPING OF THE DUTCH
ROLL OSCILLATION. John P. Campbell and
Marion 0. McKinney, Jr. October 1953. 40p.
diagrs., 4 tabs. (NACA TN 3035)

This paper presents some preliminary results of a
theoretical investigation to determine what design
features appear most important in designing modern
high-performance airplanes to have the greatest
possible inherent stability of the Dutch roll oscilla-
tion in order that the need for artificial stabilizing
devices can be minimized. These preliminary
results cover the case of fighter airplanes at sub-
sonic speeds.

NACA TN 3036

THE FLOW ABOUT A SECTION OF A FINITE-
ASPECT-RATIO NACA 0015 AIRFOIL ON A TRAN-
SONIC BUMP. Jack A. Mellenthin. October 1953.
30p. diagrs., photos., tab. (NACA TN 3036)

Pressure distributions were measured at one span-
wise station on a semispan rectangular wing model
having an NACA 0015 airfoil section and a moderate
aspect ratio. The tests were conducted on a tran-
sonic bump at Mach numbers from 0.4 to 1.06.
Pressure-distribution plots were integrated to obtain
the section lift, drag, and pitching-moment coeffi-
cients. At a fixed angle, a region developed over
the airfoil surface wherein the local Mach number
remained nearly constant as the free-stream Mach
number was increased above the critical value. This
region covered essentially the whole chord of the air-
foil at free-stream Mach numbers near unity.

NACA TN 3037

COUNTING METHODS AND EQUIPMENT FOR
MEAN-VALUE MEASUREMENTS IN TURBULENCE
RESEARCH. H. W. Liepmann and M. S. Robinson,
California Institute of Technology. October 1953.
49p. diagrs., photos. (NACA TN 3037)






NACA
RESEARCH ABSTRACTS NO.52

Methods of measuring the probability distributions
and mean values of random functions as encountered
in turbulence research were studied. Applications
to the measurement of probability distributions of the
axial velocity fluctuation u(t) and its derivative
du dt in isotropic turbulence are shown. The
assumption of independent probabilities of u(t) and
du dt was investigated and the results indicate that
the assumption is satisfied within a few percent and
that there is no evidence that the systematic differ-
ence between the microscale of turbulence measured
from zero counts and measured independently can be
traced entirely to the statistical dependence of u
and du dt. The chronological development of
apparatus is described, concluding with the 10-
channel statistical analyzer.

NACA TN 3042

HIGH-FREQUENCY PRESSURE INDICATORS FOR
AERODYNAMIC PROBLEMS. Y. T. Li,
Massachusetts Institute of Technology. November
1953. 52p. diagrs., photos., 4 tabs. (NACA
TN 3042)

Three different types of pressure indicators are
discussed. Each of these indicators has a unique
feature, but all are designed with an attempt to
combine both high-frequency response and high
resolving power into one instrument. Of the
mechanical -electrical-transducer type of pressure
indicator, the wire strain gage leads in simplicity.
The capacitance type is more versatile because it
permits the use of very high frequency carrier sys-
tems and thereby cuts down the effective interference
in the electronic system. The system utilizing the
stretching of a barium-titanate disk produces large
signals and results in compact design, but it can
only be used for dynamic measurements when tem-
perature variations are slight. Five different types
of pressure receivers were tested. The flat-
diaphragm type leads the others in simplicity, the
spherical-diaphragm type exceeds in dynamic per-
formance, -and the catenary-diaphragm type is the
one least affected by temperature change.

NACA TN 3045

ANALOGY BETWEEN MASS AND HEAT TRANSFER
WITH TURBULENT FLOW. Edmund E. Callaghan.
October 1953. 19p. diagrs. (NACA TN 3045)

An analysis of combined heat and mass transfer from
a flat plate has been made in terms of Prandtl's
simplified physical concept of the turbulent boundary
layer. The results of the analysis show that for con-
ditions of reasonably small heat and mass transfer,
the ratio of the mass- and heat-transfer coefficients
is dependent on the Reynolds number of the boundary
layer, the Prandtl number of the medium of diffusion,
and the Schmidt number of the diffusing fluid in the
medium of diffusion. For the particular case of
water evaporating into air, the ratio of mass-
transfer coefficient to heat-transfer coefficient is
found to be slightly greater than unity.



NACA TM 1360

CONCERNING THE FLOW ABOUT RING-SHAPED
COWLINGS. PART XII. TWO NEW CLASSES OF
CIRCULAR COWLS. (Uber die Stromung an


3

ringtfrmigen Verkleidungen. XII Mitteilung: Zwei
neue Klassen von Ringhaubenj. Dietrich Kiichemann
and Johanna Weber. October 1953. 72p. diagrs.,
3 tabs. (NACA TM 1360. Trans. from Zentrale
fir wissenschaftliches Berichtswesen der
Luftfahrtforschung, Berlin. UM 3111)

For application in practice for annular radiator
fairings and similar arrangements, two new classes
of circular cowls are developed by theoretical
method, and investigated in a systematic test series
regarding their behavior under various working
conditions.


NACA RM E53G03

EFFECTIVE THERMAL CONDUCTIVITIES OF
MAGNESIUM OXIDE, STAINLESS STEEL, AND
URANIUM OXIDE POWDERS IN VARIOUS GASES.
C. S. Elan and R. G. Deissler. October 1953. 18p.
diagrs., photos., tab. (NACA RM E53G03)

As a part of a general investigation of the effective
thermal conductivities of powders, tests were con-
ducted to determine the conductivity of magnesium
oxide, stainless steel, and uranium oxide powders in
various gases at temperatures between 1200 and
14550 F. Fair agreement was obtained between con-
ductivities calculated from experimental data for fine
magnesium oxide and stainless steel powders and
those calculated from a simplified analysis from a
previous investigation, although the experimental
values are somewhat higher. Runs were also made
to determine the effect of gas pressure on effective
thermal conductivity.

NACA RM E53H31

MINIMUM SPARK-IGNITION ENERGIES OF 12
PURE FUELS AT ATMOSPHERIC AND REDUCED
PRESSURE. Allen J. Metzler. October 1953. 28p.
diagrs., 5 tabs. (NACA RM E53H31)

Minimum spark-ignition energies for 12 pure fuels
were measured at reduced pressure, and the data ob-
tained were extrapolated to 1 atmosphere. The fuels
investigated included normal and cycloparaffins,
olefins, carbon disulfide, and oxygenated compounds
such as an alcohol, ether, propylene oxide, and
tetrahydropyran; these fuels were ignited at reduced
pressures by capacitance sparks of controlled dura-
tion. The minimum ignition energies obtained are
related to the pressure, the quenching distance, and
the maximum fundamental flame velocity of the fuel-
air mixture. Also, the experimental data obtained
are applied to two correlations of spark-igrunition
energies to check the data of this investigation with
that of others.

NACA RM L53I118a

FACTORS AFFECTING TRANSITION AT SUPER-
SONIC SPEEDS. K. R. Czarnecki and Archibald R.
Sinclair. November 1953. 13p. diagrs. (NACA
RM L53Il8a)

The paper surveys the available material and summa-
rizes what is known to date about boundary-layer
transition at supersonic speeds. Variables studied
include Mach number, Reynolds number, pressure
gradients, heat transfer, surface roughness, and









angle of attack. The discussion is limited to bodies
of revolution because similar reliable data for wings
is lacking.


BRITISH REPORTS


N-25356*

Aeronautical Research Council (Gt. Brit.)
WING-FUSELAGE FLUTTER OF LARGE AERO-
PLANES. W. P. Jones. 1953. 46p. diagrs., 7
tabs. (ARC R & M 2656. Formerly ARC 11,024;
0.688)

A general theoretical method is described which
takes into account a large number of degrees of free-
dom and is based on the design data for the airplane.
The problem specifically investigated is the symmet-
rical flutter of a particular aircraft. Twelve degrees
of freedom are assumed to cover pitching and trans-
lational motion of the whole airplane, flexure and
torsion of the wings, and fuselage vertical bending.
The tailplane is regarded as rigid. In the case con-
sidered, estimates indicate that the lowest critical
speed is well above the maximum design speed of the
airplane. The influence of the additional degrees of
freedom associated with movements of the control
surfaces is not considered.

N-26658*

Royal Aircraft Establishment (Gt. Brit.)
CALCULATIONS OF THE PRESSURE DISTRIBU-
TIONS AND BOUNDARY LAYER DEVELOPMENT ON
A BODY OF REVOLUTION WITH VARIOUS PARA-
BOLIC AFTERBODIES AT SUPERSONIC SPEEDS.
L. E. Fraenkel. February 1953. 56p. diagrs., 4
tabs. (RAE Aero 2482)

Detailed calculations are made of the flow over a
series of bodies at Mach numbers of 1.2, 1.4, and
1.6 and Reynolds numbers of 48 to 72 millions. The
bodies consist of a basic forebody and parallel
portion to which are added truncated parabolic after-
bodies of three different thickness ratios. Calcula-
tion of the flow over the bodies is done by the method
of characteristics. The method of Squire and Young
is used to calculate the boundary layer properties.
Calculation of the pressure distribution on the "mod-
ified" afterbodies is by Ferri's method of linearized
characteristics.

N-26687*

National Gas Turbine Establishment (Gt. Brit.)
REFLECTION OF A SMALL PRESSURE PULSE BY
DISTRIBUTED FRICTION IN ONE-DIMENSIONAL
GAS FLOW. P. W. H. Howe. July 1953. 23p.
diagrs. (NGTE Memo. M. 168)

This work is part of an investigation into the com-
bustion excited oscillations in a ram jet or reheat
system. The main interest here is how much of a
pressure pulse produced in the combustion zone
succeeds in travelling upstream past the flame
stabilizing baffle. The method used is effectively
a differentiation from the steady state-conditions
and is of general application.


NACA
RESEARCH ABSTRACTS NO.52

N-26694*

National Gas Turbine Establishment (Gt. Brit.) "-
AN IMPROVED DESIGN OF SONIC SUCTION '
PYROMETER. L. Fuller and B. Marlow. June
1953. 24p. diagrs. (NGTE Memo. M. 189)

Designs of sonic suction pyrometers for the meas-
urement of the temperature of low density, high
velocity, hot gases have been exmained and tested in
a rig that permits the comparison of two pyrometers.
one in a gas stream at approximately ground level,
atmospheric pressure and the second in the same
stream but at a pressure controllable to below 3
inches mercury absolute. Various modifications
were made and the pyrometers now advocated agree
on mean temperatures to within 50 C when one is
at 24 inches mercury absolute and the other is at 3
inches mercury absolute.


N-26842*

Aeronautical Research Council (Gt. Brit.)
NOTES ON THE TECHNIQUE EMPLOYED AT THE
R. A. E. IN LOW-SPEED WIND-TUNNEL TESTS IN
THE PERIOD 1939-1945. Edited by F. B. Bradfield.
1952. 71p. photos., diagrs., tab. (ARC R & M
2556; 11,164. Formerly RAE Aero 2222)

Very little has been recorded during the war years as
to the details of technique used in low-speed wind-
tunnel tests. The size and type of tunnel used during
this period will remain in use at firms and colleges
for some time after newer equipment is available at
research establishments, so it has been decided to
issue some record of the technique in use at the
Royal Aircraft Establihsment during the war years,
both with a view to establishing a standard technique
where it is satisfactory, and to consider weaknesses
where it has failed.


N-26843*

Aeronautical Research Council (Gt. Brit.)
THE SOLUTION OF LIFTING-PLANE PROBLEMS
BY VORTEX-LATTICE THEORY. V. M. Falkner.
1953. 30p. diagrs., 46 tabs. (ARC R & M 2591.
Formerly ARC 10,895; S & C 2153; Perf. 354)

The report describes in detail the methods by which
the principles of vortex-lattice theory, introduced in
a previous report R & M 1910, are applied to the cal-
culation of the aerodynamic loading of wings by lift-
ing plane theory. The scope of the paper is limited
to the application of these principles to symmetrical
incidence solutions and symmetrical and antisymmel-
rical wing twist solutions, for which standard solu-
tions can be treated by comparatively simple loading
functions. The effect of discontinuity of direction of
leading or trailing edge cannot be avoided even In the
simplest solutions, and it has been found necessary
to include an investigation of this problem in order to
cover the prescribed usage of the method. Special
standard functions tabulated in another report are
used to allow for the rounding off effects due to
change of direction of leading or trailing edge. The
general problem of discontinuities is under investi-
gation and will be dealt with in a later report. A
comprehensive set of solutions for a delta wing is in-
cluded in the report in order to show the convergence







NACA
RESEARCH ABSTRACTS NO.52
N-26843'

of and relation between solutions of varying com-
plexity, and to indicate which solution should be used
in order to satisfy the accuracy prescribed for any
given problem. The case of the delta wing is not
completely general, and the exposition in respect to
Induced drag and yawing moment will be completed
in a later report.

N-26847*

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF SPANWISE RIB-BOOM STIFFNESS
ON THE STRESS DISTRIBUTION NEAR A WING
CUT-OUT. E. H. Mansfield. 1952. 21p. diagrs.
(ARC R & M 2663; ARC 11,291. Formerly RAE
Structures 13)

A theoretical investigation is made into the effect of
spanwise rib-boom stiffness on the stress distribu-
tion at a cut-out in the interspar skin of a stressed
skin wing in bending. Both shear and bending stiff-
ness of the rib-boom are taken into account, and
attention is concentrated on the case in which the
rib-boom is built-in to the spar flanges. Curves are
included which determine, for any particular case, the
magnitude of the peak shear stress adjacent to the
flange, the approximate spanwise variation of this
shear stress, the proportion of load transferred by
the rib-boom to the skin and stringers, and the bend-
ing moment in the rib-boom at its points of attach-
ment to the spar flanges. By suitable design of the
rib-boom, it is possible to lower the shear stresses
adjacent to the flange with little or no increase in
structure weight.

N-26848*

Aeronautical Research Council (Gt. Brit.)
WIND-TUNNEL TESTS OF THE STALLING PROP-
ERTIES OF AN 8 PER CENT THICK SYMMETRICAL
SECTION WITH NOSE SUCTION THROUGH A
POROUS SURFACE. R. C. Pankhurst, W. G.
Raymer and A. N. Devereux. 1953. 14p. diagrs.,
tab. (ARC R & M 2666. Formerly ARC 11,496;
Perf. 441; FM 1247)

The stalling properties of an 8-percent-thick sym-
metrical airfoil with large leading-edge radius of
curvature and continuous (distributed) suction over
the nose have been tested in the 4-foot No. 2 wind
tunnel of the National Physical Laboratory. It was
found that suction postponed the stall to higher angles
of incidence by suppressing separation at the leading
edge. The suction also produced beneficial effects in
delaying transition. Moreover, it prevented the
development of boundary-layer turbulence behind a
single excresence or spanwise corrugation, provided
the suction was applied over a sufficient chordwise
extent of the airfoil surface. The quantity require-
ments are remarkably small. For example, even at
the low Reynolds number of 0.3 x 106 a quantity
coefficient CQ(Q Uc) of only 0.0036 is sufficient to
increase the lift coefficient at 150 increase by 0.6
(from 0.7 to 1.3), and it is to be expected that CQ
will become even less as the Reynolds number is
increased. It is not yet possible to estimate the
probable power requirements, because the potentiali-
ties of the best methods of porous construction are
not known.


5

N-26849*

Aeronautical Research Council (Gt. Brit.)
THE DETERMINATION OF THE NATURAL FRE-
QUENCIES OF A FULL-SCALE AIRFRAME-
ENGINE SYSTEM BY THE ADMITTANCE METHOD.
J. R. Forshaw and F. T. Mountford. 1953. 24p.
diagrs., 3 tabs. (ARC R & M 2667; ARC 11,826.
Formerly RAE Tech. Note Structures 20)

The development of the method of the measurement
of admittances and the solution of the frequency equa-
tion for a complex full-scale airframe-engine system
is given, dividing the dynamical system at the attach-
ment of the engine to the airframe, and using a force
system of equal and opposite bending moments and
shearing forces. The values of the resonance fre-
quencies obtained from the graphical solution of the
frequency equation and from the resonance test are
compared and found to be In good agreement. The
method is applicable to the matching of an engine to
an airframe by adjusting the flexibility of the mount-
ing units.

N-26853*

Aeronautical Research Council (Gt. Brit.)
HIGH-SPEED TUNNEL TESTS OF A 5 PER CENT.
CHORD DIVE-RECOVERY FLAP ON A NACA 0015
AEROFOIL. D. A. Clarke. 1953. 19p. diagrs.,
9 tabs. (ARC R & M 2689; ARC 11,743. Formerly
RAE Aero 2269)

Pressure plotting tests were made in the Royal Air-
craft Establishment high-speed tunnel on a parallel
wooden NACA 0015 wing with dive-recovery flap.
The Mach number was varied between 0.30 and 0.80,
and the Reynolds number was kept constant at
1.4 x 106. All combinations of the following were
tested: flap position 0.2c, 0.3c, 0.4c; flap angle 200,
400; Incidence 00, 40. The flap-chord/wing-chord
ratio was 0.05. The report presents a general
picture of the action of a dive-recovery flap on a
wing. The data are, however, too limited to permit
the formulation of general design recommendations.

N-26854*

Aeronautical Research Council (Gt. Brit.)
THE ROYAL AIRCRAFT ESTABLISHMENT 4 FT x
3 FT EXPERIMENTAL LOW TURBULENCE WIND
TUNNEL. PART I GENERAL FLOW CHARACTER-
ISTICS. H. B. Squire and K. G. Winter. 1953.
28p. diagrs., photos., 4 tabs. (ARC R & M 2690;
ARC 10,695; ARC 11,410. Formerly RAE Aero
2182; RAE Tech. Note Aero 1937)

The 4- by 3-foot wind tunnel was erected as a model
of larger tunnels to investigate unconventional design
features directed towards obtaining a high standard
of flow. Diffusers of 50 cone angle are used, except
for the rapid expansion through three wire-gauze
screens up to the maximum section. The contraction
ratio is 31.2:1 and nine screens are fitted in the max-
imum section. A speed control is used operating in-
dependently of the fan by means of a bypass duct.
The velocity distribution across the working-section
is constant to 1t/4 percent. The standard deviation
of the velocity with time measured over a period of
50 see is 0.03 percent. The flow in the diffusers
shows no tendency to separate and the velocity dis-
tribution approaching th# first screen is very satis-










factory. The installation of cascades with gap 'chord
ratio of 1/4 gives uniform outlet flow without apprec -
separation in the rapid expansion of the bulge, but
the flow in the contraction cone is not satisfactory.
A longer contraction would have been advantageous.
The power factor has been measured as 0.27 with all
screens fitted but could be imporved slightly if all
the leaks were sealed. The speed control is satis-
factory in operation.

N-26858*

Aeronautical Research Council (Gt. Brit.)
THE DIFFUSION OF LOAD INTO A PANEL
BOUNDED BY CONSTANT STRESS BOOMS AND A
TRANSVERSE BEAM. E. H. Mansfield. 1953. 12p.
diagrs. (ARC R & M 2729; ARC 11,885. Formerly
RAE Structures 31)

A theoretical investigation is made into the diffusion
of symmetrical, concentrated loads into a long
stiffened panel having constant stress edge members
and a transverse loading beam. Both pin-jointed and
clamped end conditions for the beam are considered.
Curves are given for determining the peak shear
stress near the boom, the variation of this shear
stress along the length of the panel, the proportion of
load transferred by the beam, and the bending
moment at the ends of the beam.


N-26859*

Aeronautical Research Council (Gt. Brit.)
AN INVESTIGATION OF THE USE OF AN
AUXILIARY SLOT TO RE-ESTABLISH LAMINAR
FLOW ON LOW-DRAG AEROFOILS. R. W.
Cumming, N. Gregory and W. S. Walker. 1953.
14p. diagrs., photos., tab. (ARC R& M 2742.
Formerly ARC 13,003; Perf. 645; FM 1424)

The use of an auxiliary slot on a laminar-flow airfoil
has been investigated to check whether laminar flow
can be reestablished by suction at the rear of the
region of deposited dirt, flies, etc. Results indicate
that in the absence of unfavorable pressure gradients,
it is possible to reestablish a laminar boundary layer
by removing a little more than the whole turbulent
layer reaching the slot, and preliminary estimates
suggest that with efficient ducting it should be possi-
ble to achieve a reduction in over-all effective drag
coefficient by this means.


N-26861*

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF COMPRESSIBLITY ON THE
ATTITUDE OF AIRCRAFT IN RECTILINEAR
FLIGHT. K. J. Lush. 1953. 7p. diagrs., tab.
(ARC R& M 2776; ARC 11,500. Formerly AAEE/
Res/234)

The attitude of aircraft (that is, the angle between the
aircraft datum and the flight path) is of considerable
importance in the aiming of certain airborne arma-
ment. An investigation was therefore made of the
effect of compressibility on the attitude of aircraft in
flight in a straight path. The application of the re-
sults of linear perturbation theory to the problem was
examined and the deductions made compared with the
results of attitude measurements on a Spitfire IX
over a wide range of altitude and airspeed.


NACA
RESEARCH ABSTRACTS NO.52

N-26862*

Aeronautical Research Council (Gt. Brit.)
THE TOEPLER SCHLIEREN APPARATUS. D. W.\
Holder and R. J. North. 1953. 13p. diagrs.,
photos., 3 tabs. (ARC R& M 2780. Formerly ARC
13.068; FM 1433)

For wind-tunnel observations it is usually necessary
that a Toepler schlieren apparatus shall have roughly
uniform sensitivity to density gradients within a cer-
tain range determined by the nature of the flow. It
is also desirable that the illumination in the image of
the flow shall be as high as possible if visual obser-
vations are to be made under the best conditions and
if photographs are to be taken with very short expo-
sures. Methods for satisfying these two require-
ments are discussed and experimental results are in-
cluded to illustrate their importance.

N-26863*

Aeronautical Research Council (Gt. Brit.)
EVAPORATION OF DROPS OF LIQUID. J. K.
Hardy. 1953. 9p. (ARC R & M 2805; ARC 10,675.
Formerly RAE Mech. Eng. 1)

An analysis has been made of the processes which
follow when a drop of liquid is subjected to a sudden
change in the condition of the air in which it is
suspended. Equations are given from which either
the temperature of the drop, or the rate at which it
will evaporate, can be calculated.


N-26886*

Royal Aircraft Establishment (Gt. Brit.)
TORSIONAL STRENGTH OF BARS IN STEEL, AND
ALUMINIUM, MAGNESIUM AND COPPER ALLOYS.
E. L. Ripley, A. J. Beard and B. A. J. McCarthy.
July 1953. 40p. diagrs., 5 tabs. (RAE Tech. Note
Structures 119)

Simple empirical relationships between the torsional
strength and tensile strength of solid and hollow bars
are derived from tests on steel, aluminum alloy,
magnesium alloy, and copper alloy bars. Both proof
and ultimate conditions of failure are considered.
The relationships are used to present design data on
the torsional strength of solid and hollow bars.


N-26887*

Royal Aircraft Establishment (Gt. Brit.)
SHEAR STRENGTH OF PINS IN STEEL, AND
ALUMINIUM, MAGNESIUM AND COPPER ALLOYS.
E. L. Ripley and A. J. Beard. July 1953. 28p.
diagrs., 5 tabs. (RAE Tech. Note Structures 120)

Simple empirical relationships between the ultimate
shear strength and ultimate tensile strength of
cylindrical pins are derived from tests on steel,
aluminum alloy, magnesium alloy, and copper alloy
bars. The relationships are used to present design
data on the shear strength of pins.






NACA
RESEARCH ABSTRACTS NO. 52 7


N-26888'

Royal Aircraft Establishment (Gt. Brit.)
LOW SPEED WIND TUNNEL MEASUREMENTS OF
THE LIFT ON A 450 SWEPT BACK HALF WING AND
CYLINDRICAL BODY. J. A. Lawford. July 1953.
40p. diagrs., 12 tabs. (RAE Tech. Note Aero 2243)

Measurements have been made of the lift, drag, and
pitching moment on a 450 sweptback half wing
mounted on a cylindrical body. Four wing aspect
ratios between 0.5 and 1.5 and two body diameters
have been tested. For the larger body the lift due to
the wing has been analyzed by pressure plotting into
contributions due to lift on the wing and to the lift
induced by the wing on the body. The measured lift
slopes are compared with values calculated by the
method of J. Weber (R.A.E. Report No. Aero 2467).
The calculation overestimates the lift slope of the
wing-body combination by approximately 10 percent.
The tests are part of an investigation to establish a
method of estimating the lift slope of this configura-
tion, which occurs when a swept fin is carried on a
body of circular section (for example, a jetpipe) and
when a single store is carried on a wing tip.


N-26889*

Aeronautical Research Council (Gt. Brit.)
INVESTIGATION OF THE INSTABILITY OF A
MOVING LIQUID FILM. H. B. Squire. January 25,
1952. 10p. photos., diagrs., tab. (ARC 14,586;
CF 213)

The stability of a thin layer of liquid moving in still
air is studied theoretically with the object of throw-
ing light on the break-up of films during atomization.
It is found that instability occurs if W = T/plU2h< 1
and that the wave length for maximum growth factor,
for W 1, is X = 4n7T/p2U2 where p1 is the
liquid density, p2 is the air density, U is the film
velocity, 2h is the film thickness, and T is the
surface tension of the liquid. Comparison with ex-
perimental data shows fair agreement with the
observed wave lengths.


N-26958*

Royal Aircraft Establishment (Gt. Brit.)
A TECHNIQUE FOR STUDYING THE BEHAVIOUR OF
CINE CAMERA MECHANISMS UNDER DYNAMIC
CONDITIONS. G. L. Davies. July 1953. 10p.
diagrs., photos. (RAE Tech. Note GW 261)

In the design of high-speed cine camera mechanisms,
it is not always possible to predict the behavior of
the film under dynamic conditions. A simple tech-
nique has been developed for measuring running
speed, register, film acceleration, shutter timing,
and other parameters. The camera under test has
its normal shutter replaced by a "negative" shutter,
and it is then used to photograph a cathode ray
oscilloscope which has a linear saw tooth wave form
applied to its X deflection plates. The wave shape of
the resulting traces on the film can easily be in-
terpreted to give the required information.


NACA-Langley 11-17-53 4M


UNPUBLISHED PAPERS

N-25367*

THE MODERN DEVELOPMENT OF THE THEORY
OF PROPELLERS. (Le developpement moderne de
la theorie de I'helice). Raymond Siestrunck.
September 1953. ii, 137p. diagrs. (Trans. from
Institute de Mecanique de la Faculte des Sciences de
Paris, 1947)

The empirical basis of propeller vortex theory is
examined; theories already known are arranged and
summarized. The general form of the vortex
surfaces with respect to the propeller is specified
and evaluation of the field of velocities induced by
these vortices is undertaken. The idea is given that
the functioning of the propeller depends on the state
of the velocities of the surrounding fluid. The
problem of determining the distribution of the circu-
lations which induce the velocities along a blade is
studied in detail as are the problems of its adapta-
tions for practical use.



MISCELLANEOUS


NACA TN 2135

Errata No. 1 on "THE CALCULATION OF DOWN-
WASH BEHIND WINGS OF ARBITRARY PLAN
FORM AT SUPERSONIC SPEEDS. John C. Martin.
July 1950.



NACA TN 2590

Errata No. 1 on "CALCULATIONS ON THE FORCES
AND MOMENTS FOR AN OSCILLATING WING-
AILERON COMBINATION IN TWO-DIMENSIONAL
POTENTIAL FLOW AT SONIC SPEED." Herbert C.
Nelson and Julian H. Berman. January 1952.





UNIVERSITY OF FLORIDA


3 1262 08153 2649




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