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:
AA00009235:00041

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National Advisory Committee for ,ronautica



Research Abstract \
3.78 YfBRUARrYAiPLS e


CURRENT NACA REPORTS

NACA Rept. 1165

UNSTEADY OBLIQUE INTERACTION OF A SHOCK
WAVE WITH A PLANE DISTURBANCE. Franklin K.
Moore. 1954. ii, 21p. diagrs. (NACA Rept. 1165.
Formerly TN 2879)

Analysis is made of the flow field produced by oblique
impingement of weak plane disturbances of arbitrary
profile on a plane normal shock. Three types of dis-
turbance are considered: (a) Sound wave propagating
in the gas at rest into which the shock moves. The
sound wave refracts either as a simple isentropic
sound wave or an attenuating isentropic pressure
wave, depending on the angle between the shock and
the incident sound wave. A stationary vorticity wave
of constant pressure appears behind the shock. (b)
Sound wave overtaking the shock from behind. The
sound wave reflects as a sound wave. and a stationary
vorticity wave is produced. (c) An incompressible
vorticity wave stationary in the gas ahead of the shock.
The incident wave refracts as a stationary vorticity
wave, and either a sound wave or attenuating pres-
sure wave is also produced. Computations are pr .~
sented for the first two types of incident wave, er -
the range of incidence angles, for shock Mach
*numbers of 1, 1.5, and. .



NACA Rept. 1167

METHOD FOR CALCULATING THE ROLLING A
YAWING MOMENTS DUE TO ROLLING FOR UN- "-. 4
SWEPT WINGS WITH OR WITHOUT FLAPS OR
AILERONS BY USE OF NONLINEAR SECTION LIFT
DATA. Albert P. Martina. 1954. ii, 16p. diagrs.,
11tabs. (NACA Rept. 1167. Formerly TN 2937)

The methods of NACA Reports 865 and 1090 have
been applied to the calculation of the rolling- and
yawing-moment coefficients due to rolling for un-
swept wings with or without flaps or ailerons. The
methods allow the use of nonlinear section lift data
together with lifting-line theory. Two calculated
examples are presented in simplified computing
forms in order to illustrate the procedures involved.



NACA Rept. 1168

SECONDARY FLOWS AND BOUNDARY-LAYER
ACCUMULATIONS IN TURBINE NOZZLES. Harold
E. Rohlik, Milton G. Kofskey, Hubert W. Alien and
Howard Z. Herzig. 1954. ii, 32p. diagrs., photos.,
3 tabs. (NACA Rept. 1168. Formerly TN 2871;
TN 2909; TN 2989)


*AVAILABLE ON LOAN ONLY.


ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 15t1 H ST., NW.,
THE REPORT TITLE AND AUTHOR.
S 2q. 13092-
i 77.S" -


The results of detailed measurements of the high-
speed flows in three turbine nozzle configurations
are interpreted, with the aid of low-speed flow-
visualization experiments, in order to establish the
sources and patterns of the secondary flows there.
The secondary-flow mechanism causes local accumu-
lation of high-loss material by transportingooundary-
layer fluids across the channels on the walls, as
well as radially in the blade wakes and boundary
layers. The size and location of these boundary-
layer accumulations depend largely on the blade
velocity profiles, not on the radial distribution of
circulation along the blades.



NACA Rept. 1170

BEHAVIOR OF MATERIALS UNDER CONDITIONS
OF THERMAL STRESS. S.S. Manson. 1954. ti,
34p. diagrs., photos., 6 tabs. (NACA Rept. 1170.
Formperly TN 2933)

E review is presented of available information on the
'ior of brittle and ductile materials under condi-
Sritj f thermal stress and thermal shock. For
br flde materials, simple formulas relating physical
t tes to thermal-shock resistance are derived
San d. sr to determine the relative significance of
tworI les currently in use for rating materials.
T i ortance of simulating operating conditions in
tr I-shock testing is deduced from the formula
- ndj' experimentally illustrated by showing that
S could be both inferior or superior to A1203 in
thermal shock depending on the testing conditions.
For ductile materials, thermal-shock resistance
depends upon the complex interrelation among
several metallurgical variables which seriously
affect strength and ductility. These variables are
briefly discussed and illustrated from literature
sources. The importance of simulating operating
conditions in tests for rating ductile materials is
especially to be emphasized because of the
importance of testing conditions in metallurgy. A
number of practical methods that have been used to
minimize the deleterious effects of thermal stress
and thermal shock are outlined.



NACA TN 3347

A WIND-TUNNEL TEST TECHNIQUE FOR
MEASURING THE DYNAMIC ROTARY STABILITY
DERIVATIVES INCLUDING THE CROSS DERIVA-
TIVES AT HIGH MACH NUMBERS. Benjamin H.
Beam. January 1955. 35p. diagrs., photos.
(NACA TN 3347)


WASHINGTON 25, D.C., CITING CODE NUMBER ABOVE EACH TITLE,


N(


_ _










An experimental technique for measuring dynamic
stability derivatives of a model airplane in a wind
tunnel is described. A single-degree-of-freedom,
feedback-controlled, forced-oscillation system was
used to measure the stability derivatives which are
important in estimating the lateral and longitudinal
oscillatory motions of a rigid airplane. Some
representative experimental data are included.




NACA TN 3350

THE LINEARIZED EQUATIONS OF MOTION
UNDERLYING THE DYNAMIC STABILITY OF AIR-
CRAFT, SPINNING PROJECTILES, AND SYM-
METRICAL MISSILES. A. C. Charters. January
1955. 102p. diagrs. (NACA TN 3350)

Linearized equations of motion are derived for air-
craft with mirror symmetry and for spinning pro-
jectiles with both rotational and mirror symmetry.
The customary set of dynamic stability derivatives
is extended to include the aerodynamic effects of
spin. The aircraft and projectile cases are treated
alike and the consequences of symmetry and spin are
shown by a comparison of the equations for the two
cases. Dynamic stability conditions are derived.
The equations for spinning projectiles are applied
to the analysis of flight test data from the aero-
dynamics range.




NACA TN 3358

GUST-LOAD AND AIRSPEED DATA FROM ONE
TYPE OF FOUR-ENGINE AIRPLANE ON FIVE
ROUTES FROM 1947 TO 1954. Walter G. Walker.
January 1955. 28p. diagrs., 4 tabs. (NACA
TN 3358)

This paper presents the results of an analysis of
approximately 100,000 hours of V-G data from one
type of four-engine civil transport airplane to deter-
mine the magnitude and frequency of occurrence of
the gust loads and gusts. The normal accelerations
differed by approximately 15 percent and the derived
gust velocities by about 18 percent for the five oper-
ations investigated. The gust loads of the present
operations were less than the loads experienced by
other four-engine civil transports previously inves-
tigated, but the differences are not significant. The
present data indicated only small differences due to
seasonal effects and different operational utilization.




NACA TN 3363

LOW-SPEED WIND-TUNNEL INVESTIGATION OF A
TRIANGULAR SWEPTBACK AIR INLET IN THE
ROOT OF A 450 SWEPTBACK WING. Arvid L.
Keith, Jr. and Jack Schiff. January 1955. 65p.
diagrs., photos., 5 tabs. (NACA TN 3363. Formerly
RM L50101)

Results of a low-speed study of a 450 sweptback
wing-root air-inlet configuration believed suitable
for transonic-speed airplanes are presented. The


NACA
RESEARCH ABSTRACTS NO. 78


inlet -configuration lilt and drag characteristics are
compared with those of a basic model. Boundary-
layer growth along the fuselage nose, inlet total-
pressure recoveries, and static-pressure distribu-
tions over the inlet and wing surfaces are presented
for wide ranges of inlet-velocity ratio and angle of
attack.




NACA TN 3386

SOME CONSIDERATIONS ON TWO-DIMENSIONAL
THIN AIRFOILS DEFORMING IN SUPERSONIC
FLOW. Eugene Migotsky. January 1955. 36p.
diagrs. (NACA TN 3386)

The aerodynamic characteristics of indicially cam-
bered two-dimensional airfoils in supersonic flow
are determined theoretically. The power required
to sustain a general time-varying chordwise defor-
mation is determined. Stability boundaries are
presented for the harmorucally oscillating parabolic
mode. The thickness distribution of a beam having
a parabolic fundamental bending mode, in vacuo, is
determined.



NACA TN 3388

A FIBROUS-GLASS COMPACT AS A PERMEABLE
MATERIAL FOR BOUNDARY-LAYER-CONTROL
APPLICATIONS USING AREA SUCTION. Robert E.
Dannenberg. James A. Weiberg and Bruno J.
Gambucci. January 1955. 20p. diagrs., photos.,
2 tabs. (NACA TN 3388)

The resistance to air flow of fibrous-glass compacts
suitable for boundary-layer control was measured.
The flow resistance is related to the thickness and
density of the fibrous-glass compacts. Constant
thickness compacts with specified permeability
distributions were fabricated and tested.




BRITISH REPORTS




N-34689*

Nat. Gas Turbine Establishment (Gt. Brit.)
THE DEVELOPMENT OF AN ACCURATE SPEED
MEASURING DEVICE (THE TACHRONOMETER).
H. Shaw and W. E. Wilcox. May 1954. 12p. diagrs.,
photos. (NGTE Memo. M. 216)

An electrical system has been developed which will
control a stopwatch to time a preselected number of
revolutions of a shaft. Special attention has been
given to the removal of sources of error, and to the
provision for testing the operation of the instrument
in situ. An accuracy of 0.2 percent can be obtained
for mean speeds measured over a period of at least
50 seconds. The instrument is used in conjunction
with an electric contact-maker driven through a
reduction gear from the shaft of which the speed is
to be measured.






NACA
RESEARCH ABSTRACTS NO. 78

N-34690*

Nat. Gas Turbine Establishment (Gt. Brit.)
ON RENDERING ENGINE EXHAUST GASES NON-
EXPLOSIBLE. B. P. Mullins and J. M Hawkins.
October 1954. 34p. diagrs. (NGTE Memo. M. 225)

A quantitative assessment of explosion hazards in
combustion system exhaust gases has been made,
and a criterion of nonexplosibility defined. Calcula-
tions have been made of: (1) the limiting quantities
of induced air and injected water required to render
exhaust gases of fuel burrung rigs nonexplosible at
all times; (2) quantities of air and water necessary
to reduce the exhaust duct gas temperature to any
desired level; (3) limits of explosibility m the static
pressure range 0.1 to 1.0 atm when some fraction of
the fuel supply is burned and the gas subsequently
cooled to below 1000 C; and (4) limiting safe fuel:
air ratios at various initial mixture temperatures
and pressures in the absence of injected water.


N -3469 1'

Royal Aircraft Establishment (Gt. Brit.)
AN IMPROVED PRESSURE TRANSDUCER EMPLOY-
ING AN ELLIPTICAL TUBE AND STRAIN GAUGES.
D. S. Dean. July 1954. 26p. diagrs., photos. (RAE
Tech. Note RPD 106)

A pressure transducer is described which employs
bonded wire strain gages to measure circumferential
strain in a short elliptical tube. The transducer is
designed for high sensitivity (double that possible
with the usual circular tube type), robustness, ac-
curacy, good temperature compensation and ability
to withstand overloads. Graphs are included from
which all the data may be obtained to enable trans-
ducers to be constructed for a wide variety of pres-
sure ranges and to suit various electrical circuits.
The transducer is covered by U. K. provisional
patent Specification No. 21479 53.


N-34694*

Royal Aircraft Establishment (Gt. Brit.)
EVALUATING THE SUPERSONIC DRAG INTEGRAL
USING FOURIER SERIES. N. A. Routledge.
September 1954. 16p. (RAE Tech. Note MS 17)

The expression


Sj J "x) S"(y) logI x y-L dx dy,
0 0

which is related to the drag of a body travelling at
supersonic speeds, is evaluated for arbitrary
functions S(x) such that S'(o) = S'(l) = 0 by
fitting Fourier Sine Series. Some general remarks
are made on the utility of the method of approach
that is adopted.



N-34721*

Aeronautical Research Council (Gt.Brit.)
A NOTE ON THE DEVELOPMENT OF SENSITIVE
PRESSURE OPERATED WATER CONTACTS FOR
USE ON SEAPLANES. R. Parker. 1954. lip.
diagrs., photos., tab. (ARC CP 176)


3


A pressure operated water contact has been devel-
oped, suitable for indicating the instants of take-off
and touch-down for a seaplane hull. Flight tests
have shown that the instrument is accurate In opera-
tion and sufficiently robust for normal flight test use



N-34772*

Aeronautical Research Council (Gt.Brit.l
A SURVEY OF SCALE EFFECTS ON THE HYDRO-
DYNAMIC TESTING OF SEAPLANE MODELS.
R. Parker. 1954. 29p. diagrs., photos., tab.
(ARC CP 179)

A general survey is made of all the factors where
true dynamic similarity cannot be achieved in model
tests of seaplane hulls, and the likely effects on test
results are discussed with reference to towing tank
models and medium size research aircraft. In resis-
tance tests the correction for Reynolds number effects
requires more investigation. Pressure effects are
likely to affect the breakaway of flow at small dis-
continuities such as extreme fairings with resultant
errors in both stability and resistance test results.
More accurate and systematic full scale data than at
present is available is needed before methods of
allowing for this can be satisfactorily developed.



N-34777'

Aeronautical Research Council (Gt. Brit.)
PUBLISHED REPORTS AND MEMORANDA FOR THE
AERONAUTICAL RESEARCH COUNCIL. 1954.
7p. (ARC R M 2650)




N-34778'

Aeronautical Research Council (Gt.Brit.)
OPTICAL CONSIDERATIONS AND LIMITATIONS OF
THE SCHLIEREN METHOD. G. S. Speak and D. J.
Walters. 1954. 25p. diagrs., 6 tabs. (ARC
R u M 2859; 13,066. Formerly RAE Tech. Note
LAP 968)

The elementary principles of the schlieren method
are described, with reference to an ideal basic sys-
tem. The experimental procedure in setting up the
system is covered from the same aspect. It is con-
cluded that the schlieren method may be used quali-
tatively at extremely high sensitivities with satisfac-
tory results, but is not suitable for quantitative work
where small pressure or density changes are in-
volved. A secondary conclusion is that the twin-
mirror system is in general the best for overall
ease of interpretation of results, though local con-
siderations may modify this choice.



N-35208'

Aeronautical Research Council (Gt.Brit.)
AUTHOR INDEX TO THE REPORTS AND MEMO-
RANDA AND CURRENT PAPERS OF THE AERO-
NAUTICAL RESEARCH COUNCIL. (1909-
January 1954) 190p. (ARC R & M 2570)




UNIVERSITY OF FLORIDA
m lltl l IIIII II ~lII1II~ l tll ~I /1 II


4 MIII1111 IIIII
3 1262


MISCELLANEOUS


NACA Rept. 1165

Errata on "UNSTEADY OBLIQUE INTERACTION OF
A SHOCK WAVE WITH A PLANE DISTURBANCE"
Franklin K. Moore. 1954.




N-35053'

FIFTY YEARS OF BOUNDARY LAYER THEORY
AND EXPERIMENT. Hugh L. Dryden. (Presented
at meeting of Fluid Dynamics Div., Amer. Physical
Soc., Hotel Chamberlin, Old Point Comfort, Va.,
Nov. 23, 1954). 17p.

This paper summarizes boundary-layer theory and
experiment since Prandtl's boundary-layer theory
was presented 50 years ago. It tells the story of
the development of the new concept, its slow ac-
ceptance and growth, and its spread from group to
group within the country of origin and its diffusion to
other countries of the world.






UNPUBLISHED PAPERS







N-35089*

THE S. N. E. C. M. A. TRANSONIC WIND TUNNEL.
(La Soufflerie Transsonique de la S. N. E. C. M. A.).
O. Frenzl. November 1954. 30p. diagrs., photos.
(Trans. from Docaero, no.23, Sept., 1953, p. 1-18)

The new S. N. E. C. M. A. wind tunnel is operated by
a vapor ejector. After a discussion of its principle,
the report describes the air reheater with gravel
designed to prevent condensation, the antiturbulence


081


I 111111111 $\l1MlII NACA
153299 5 RESEARCH ABSTRACTS NO.78

grid. the test section with semiguided jet of 400 x
600 rectangular cross section, of 2.70 m in length.
the control valves, the ejector, the diffuser, and the
steam accumulator. The three-component wind-
tunnel balance witn six pressure cells, the schlieren
apparatus, the Mach meter, and multiple pressure
gage are briefly discussed, and the more particular-
ly satisfactory results of the first tests are reported.



N-35126'

Cornell Uriversity, Graduate School of Aeronautical
Engineering. ON THE GROWTH OF PLASTIC
DEFORMATION IN A BAR SUBMITTED TO LONGI-
TUDINAL IMPACT. C. Riparbelli. January 1953.
41p. diagrs. ICornell University, Graduate School
of Aeronautical Engineering)

The deformation of a semi-infinite bar submitted to
end impact is computed step-by-step, each step
corresponding to one element of a bar and to one
time interval. At every step, two phases are con-
sidered: (a) propagation of the deformation wave as
if it were elastic. (b) increment of deformation due
to creep.



N -35237

Cornell University, Graduate School of Aeronautical
Engineering. DISTRIBUTION OF STRESS, STRAIN,
VELOCITY IN A BAR OF NON-UNIFORM CROSS-
SECTION UNDERGOING LONGITUDINAL IMPACT
IN THE PLASTIC RANGE. Carlo Riparbelli.
May 15, 1954. 77p. diagrs.. photos. (Cornell
University, Graduate School of Aeronautical
Engineering)

A method for computing the distribution of stress,
strain, and velocity in a bar of nonuniform cross
section in the plastic domain is offered. This is
based on the separate computation of two simultaneous
phenomena, namely, propagation of elastic strain
component, taking into account the taper effect, and
plastic flow. The computation is performed step-by-
step and the above said two phases are considered at
each step. The distribution of plastic strain is thus
predicted under the assumption of its being irrevers-
ible. Comparison is offered with distributions of
plastic strain from experiments on copper bars of
nonuniform cross section.


NACA-Langley 2-17-55 4M




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