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 )
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serial   ( sobekcm )
federal government publication   ( marcgt )
abstract or summary   ( marcgt )

Notes

Statement of Responsibility:
National Advisory Committee for Aeronautics.
Dates or Sequential Designation:
Abstracts no. 1 (June 15, 1951)-no. 93 (Nov. 30, 1955).

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:
AA00009235:00027

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Full Text

NACA RM 54D19

PRELIMINARY MEASUREMENTS OF TURBULENCE
AND TEMPERATURE FLUCTUATIONS BEHIND A
HEATED GRID. A. L. Kistler, V. O'Brien and S.
Corrsin, Johns Hopkins University. June 1954.
24p. diagrs., photo. (NACA RM 54D)19)

Preliminary measurements have been made of
velocity and temperature fluctuations in the flow be-
hind a heated grid in a uniform air stream. Tem-
perature correlation shows a reasonable degree of
isotropy, and the temperature fluctuations die out at
large distances more slowly than the turbulence, as
has been predicted theoretically under some strongly
simplifying postulates.

NACA TN 3072

A THEORETICAL INVESTIGATION OF THE AERO-
DYNAMICS OF WING-TAIL COMBINATIONS PER-
FORMING TIME-DEPENDENT MOTIONS AT
SUPERSONIC SPEEDS. John C. Martin, Margaret
S. Diederich and Percy J. Bobbtitt. May 1954. 226p.
~~diagra tab. (NACA TN 3072)

A 11s.reticrlal Investigation is presented on the con-
i rfllestIr.n ...fth horizontal tail to the lift and plrconup~
nlmort" due ~to gle of attack, a constant rate of
..):tch, and a conlstant vertical acceleration. Numer-
(al vallues ... the aerodynamic coefficients associated
wlth these n...ts...ns are presented for a number of
/(w:,r'dlnlensl..nalI wing-tail combinations, a triangular-
muu;-triangulalr-tail combination, and a number of
ft-calanular -wing-triangular-tail combinations. A
.method of treating unsteady aerodynamics based on an
infinite series of stabasel .ier staat srIr of successively
higher order is presented. Methods for calculat-
ing the flow fields behind winlgs with a constant
vertical acceleration are developed. Calculated re-
sults are presented for the upwash behind two-
dimensional wings and for certain regions behind
triangular and rectangular wings for a constant rate
of pitch and for a constant rate of vertical accelera-
tion.

NACA TN 3150

METHOD FOR RAPID DETERMINATION OF PRES-
SURE CHANGE FOR ONE-DIMENSIONAL FLOW
WITH HEAT TRANSFER, FRICTION, ROTATION,
AND AREA CHANGE. James E. Hubbartt, Henry O.
Slone and Vernon L. Arne. June 1954, 22p. diagrs.,
2 tabs. (NACA TN 3150)

An approximate method for rapid determination of
the pressure change for subsonic flow of a com-
pressible fluid under the simultaneous action of heat
transfer, friction, rotation, and area change is de-
veloped. In the development of this method, the
momentum equation was approximated and re-


CURRENT NACA REPORTS

NACA Rept. 1132

LAMIINAR BOUNDARY LAYER ON CONE IN SUPER-
SONIC FLOW AT LARGE ANGLE OF ATTACK.
Franklin K. Moore. Appendix B: NUMERICAL
SOLUTION OF DIFFERENTIAL EQUATIONS. Lynn
Albers. 1953. ii,13p. diagrs. (NACA Rept. 1132.
Formerly TN 2844)

The laminar boundary-layer flow about a cone at
large angles of attack to a supersonic stream has
been analyzed in the plane of symmetry. At the
bottom of the cone, velocity profiles were obtained
showing the expected tendency of the boundary layer
to become thinner on the under side of the cone as
the angle of attack is increased. At the top of the
cone, the analysis failed to yield unique solutions,
except for small angle of attack. Beyond a certain
critical angle of attack, boundary-layer flow does
not exist in the plane of symmetry, thus indicating
separation.

'
NACA Rept. 1139 -.

CHARTS AND APPROXIMATE FORMULAS i~HE
ESTIMATION OF AEROELASTIC EFFECT I jTHE
LATERAL CONTROL OF SWEPT AND UN UE'q!}[ I
WINGS. Kenneth A. Foss and Franklin H'.
Diederich. 1953. iii, 25p. diagrs.. 2jtabs
(NACA Rept. 1139. Formerly TN 2747)

Charts and approximate formulas are presented Mr +4
the estimation of static aeroelastic effects on the -
spanwise lift distribution, rolling-moment coeffi-
cient, and rate of roll due to aileron deflections on
swept and unswept wings at subsonic and supersonic
speeds. Some design considerations brought out by
the results of this paper are discussed.


NACA Rept. 1148

A SPECIAL INVESTIGATION TO DEVELOP A
GENERAL METHOD FOR THREE-DIMENSIONAL
PHOTOELASTIC STRESS ANALYSIS. M. M. Frocht
and R. Guernsey, Jr., Illinois Institute of Technology*
1953. ii, 17p. diagrs., photos., 3 tabs. (NACA
Rept. 1148. Formerly TN 2822)

The method of strain measurement after annealing is
reviewed and found to be unsatisfactory for the
materials available in this country. A new, general
method is described for the photoelastic determina-
tion of the principal stresses at any point of a gen-
eral body subjected to arbitrary loads. The method
has been applied to a sphere subjected to diameltral
compressive loads. The results show possibilities
of high accuracy.


*AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512n H ST., NW.,
THE REPORT TITLE AND AUTHOR.


WASHINGTON Is. D. C., CITING CODE NUMBER ABOVE EACH TITLE;


National Advisory Committee For Ae~ronautics


Research Abstracts


libE_ 22 .1.~954-


NO. 65










arranged for a convenient solution employing charts.
This report presents the analysis involved in simpli-
fying the momentum equation, the charts necessary
for obtaining particular solutions, and comparisons
with more rigorous numerical solutions with condi-
tions typical for air-cooled turbine blades.



NACA TN 3174

INFLUENCE OF AIRFOIL TRAILING-EDGE ANGLE
AND TRAILING-EDGE-THICKNESS VARIATION ON
THE EFFECTIVENESS OF A PLAIN FLAP AT HIGH
SUBSONIC MACH NUMBERS. Albert D. Hemenover
and Donald J. Graham. June 1954. 101p. diagrs.,
photos., 5 tabs. (NACA TN 3174. Formerly
RM A51C12a)

The effects of variation of trailing-edge angle and
trailing-edge thickness on the characteristics of a
10-percent-chord thick NACA airfoil section with a
25-percent-chord plain flap are appraised from
wind-tunnel tests at Mach numbers from 0.3 to 0.9
and Reynolds numbers varying correspondingly from
1 to 2 million. The airfoil trailing-edge angle was
varied from 180 to 60, and the trailing-edge thick-
ness from zero to the thickness at the flap hinge line.



NACA TN 3207

ROLE OF NICKEL DIP IN ENAMELING OF SHEET
STEEL. D. G. Moore, J. W. Pitts and W. N'
Harrison, National Bureau of Standards. June 1954.
27p. diagrs., photos., 8 tabs. (NACA TN 3207)

An investigation was made of the effects of the firing
time and the weight of the nickel deposited from the
nickel-dip solution on the adherence developed by a
cobalt-free and a cobalt-bearing ground-coat enamel
on both enameling iron and a titanium-bearing low-
carbon steel.



NACA TN 3215

TESTS OF BONDED AND RIVETED SHEET-
STRINGER PANELS. Leonard Mordfin and I. E.
Wilks, National Bureau of Standards. June 1954.
45p. diagrs., photos., 5 tabs. (NACA TN 3215)

Bending and compression tests were performed on
21 sheet-stringer panels of 75S-T6 aluminum alloy
having alclad sheets nominally 0.051 inch thick and
stringers nominally 3-1/2 inches apart. Nine panels
had I-stringers bonded to the sheets with Araldite
Type I adhesive, three panels had I-stringers bonded
to the sheets with Metlbond adhesive, and nine panels
had Z-stringers riveted to the sheets. All tests
were carried to failure and strain and deformation
measurements were made. The test results did not
indicate any great superiority of one type of con-
struction over the other but rather that the choice in
any given case would depend upon the particular de-
signs being compared. They also showed that the


riee co="stu t ;-n and h hd cla ag a n
always the governing factor in the strength of bonded
panels.


NACA
RESEARCH


ABSTRACTS NO. 65


BIRITISH REPORTS




N-30793*

Royal Aircraft Establishment (Gt. Brit.)
A HIGH SPEED ELECTRO-MECHANICAL MULTI-
PLIER. C. A. A. Wass and D. W. Allen. February
1954. 13p. diagrs., photos. (RAE Tech.Note GW 300)

Requirements exist in simulator work for a simple
high-speed squaring and multiplying unit of moderate
accuracy. This note describes a simple electro-
mechanical multiplier using a type E.4 relay as a
position servo driving a potentiometer. The unit will
operate on frequencies up to 15 eps without intro-
ducing any appreciable phase shift. Accuracy de-
pends primarily on the quality of the potentiometers
used but maximum errors of 1 percent of the maxi-
mum output appear possible.





N-30835*

Royal Aircraft Establishment (Gt. Brit.)
A NOTE ON THE USE OF STRAIN GAUGES IN WIND
TUNNEL BALANCES. J. R. Anderson. January
1954. 28p. diagrs., photos. (RAE Tech.Note
Aero 2290)

This note is substantially that presented at the
NATO/AGARD wind tunnel and model testing panel
on September 3, 1953. It reviews briefly some of
the experience obtained in the employment of wind-
tunnel balances using bonded, resistance type elec-
trical strain gages, in the smaller high-speed wind
tunnels of Aerodynamics Department. The effects
of temperature on the strain gages are discussed and
shown to be a major limitation in their use. A self -
balancing potentiometer instrument, specially de-
signed and developed for use with wind-tunnel strain-
gage balances, is also described briefly.





N-30837t

Marine Aircraft Experimental Establishment
(Gt. Brit. ) THE DEVELOPMENT OF A MINIATURE
PRESSURE PICK-UP. J. K. Friswell. March 1954.
10p. diagrs., photos. (MAEE F/TN/3)

Details are given of the construction of a miniature
pressure pickup of the diaphragm and strain-gaged
cantilever type. The unit is cylindrical, with over-
all diameter 1 inch and overall length 2 inches, its
weight being 1. 9 ounces. It is designed to operate

in te range 020 esiand bilrcodee r po tie

easy to service.







NACA
RESEARCH ABSTRACTS NO. 65



N-30839

Marine Aircraft Experimental Establishment
(Gt. Brit. ) ASSESSMENT OF REVERSIBLE PITCH
PROPELLER TRAIL INSTALLATION SOLENT
N. J.201. J. Taylor. March 1954. 23p. diagrs.,
photos., 2 tabs. (MAEE F. TX '2)

An assessment was required on the operation of the
reversible pitch propeller trial installation, on the
Solent N. J. 201, as part of a general research pro-
gram for seaplanes. Results are given with the
reversing pitch propellers fitted to the inboard
engines in tests on low speed maneuvering, partic-
ularly in buoy approaches, and when used as braking
units during landing. Some brief turning circle
measurements have also been done.





N-30843*

Aeronautical Research Council (Gt. Brit. )
THE PERFORMANCE OF THE 108 COMPRESSOR
FITTED WITH LOW STAGGER FREE VORTEX
BLADING. D. V. Foster. 1954, 36p. diagrs.,
photos. (ARC CP 144)

A large three-stage compressor is described which
is designed for detail three-dimensional flow inves-
tigations. Particular attention has been paid to the
accuracy of measurement on the rig and it is shown
that, the main errors are due to the unsteady nature
of the flow and to speed fluctuations. Test charac-
teristics of the first set of low stagger free vortex
blading are presented and compared with various
theoretical performance calculations. A description
of the surging behavior of the compressor and the
effect of blade position upon the measured static
pressure are given.






N-30844+

Aeronautical Research Council (Gt. Brit. )
ZERO LIFT DRAG MEASUREMENTS ON SWEPT
WINGS AT TRANSONIC AN SUPERSONIC SPEEDS
USING THE GROUND-LAUNCHED ROCKET-
BOOSTED MODEL TECHNIQUE. T. Lawrence and
C. Kell. 1954. 35p. diagrs., photo., tab. (ARC
CP 145)

This is mainly a documentary record of drag meas-
urements on 14 swept wings varying in plan form
from deltas to swept untapered wings from 4 percent
to 10 percent thick. Results are compared with
theory for wrings of double-wedge section; an attempt
is made to check the validity of supersonic similarity
laws. Three primary conclusions are drawn; at
supersonic speeds, wave drag of a given wing varies
as the square of thickness ratio; supersonic similar-
ity law allows drag of "similar" wings to be com-
pared; drag of round-nose wings perhaps may be
estimated from calculations of drag of double-wedge
wings.


N-30845"

Aeronautical Research Counlcil (Gt. Brit. )
VIBRATION AND FLUTTER OF AIRCRAFT
AERIALS. W. H. Johnson. 1954. 15p. diagrs.,
photos. (ARC CP 14tI?

To swl failures of blade and whip aircraft aerials
have been investigated and it is shown that stalling
flutter and mechanically excited vibration hav~e both
contributed in large measure to the failures. All the
aerial types involved possess considerable 111 .t. Inl,
and very low internal damping. It is shown that the
introduction of damping into the ne-unting, of the
aerials has a very beneficial effect on their behavior
as regards both flutter and mechanically excited
vibration. Methods are described for preventing
failure from either cause.






N-30846'

Aeronautical Research Council (Gt. Brit. )
FACTORS INFLUENCING THE OPTIMUM AERO-
DYNAMIC DESIGN OF COOLED TURBINES.
G. F. C. Rogers. 1954. 27p. diagrs. (ARC
CP 155)

Owing to the losses in performance which increase
with the rate of heat extraction required for turbine
cooling, it is desirable to know what range of aero-
dynamic designs is associated with low values of this
quantity. Different aerodynamic designs of turbine,
all passing the same mass flow and having approxi-
niately the same disc and blade stresses, have been
compared on the basis of the ratio of heat extraction
rate to work output. It is found that high values of
flow coefficient are beneficial in this respect, and
that impulse turbines have a slight advantage over
reaction turbines.






N-30847*

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF ROLLING ON FIN-AND)-RUDDER
LOADS IN YAWING MANOEUVRES. D. R. Puttock.
1954. 31p. diagrs., 3tabs. (ARC CP 153)

Exact solutions are derived for angle of sidestip and
fin-and-rudder loads for an aircraft performing two
yawing maneuvers induced by the rudder. Angles of
sideslip and fin-and-rudder loads are then calculated
for three selected aircraft and compared with results
obtained by a simplified method in which rolling
motion is neglected. Further calculations are made
using a modified method in which the coefficients of
the response formulas of the simplified method have
been adjusted to take some account of rolling. The
analysis shows that errors of 20 percent may be
incurred if rolling is neglected in the estimation of
fin-and-rudder loads for aircraft with swept and
delta wings.








R ARCH ABSTRACTS NO. 65


N-30895'


Royal Aircraft Establishmlent (Gt. Brit.)
SECONDARY FLOW IN A BOUNDARY LAYER. L.
Sowerby. March 1954. 21p. diagrs., tab. (RAE
Aero 2512)

The laminar boundary layer on a flat plate is con-
sidered. It is shown that a secondary flow in the
boundary layer is a consequence of curvature in the
streamlines of the mainstream flow (in planes paral-
lel to the plate). The special case in which the
streamlines form a family of translates is discussed
and the equations of possible families of streamlines
are determined. The boundary-layer equations are
solved in the case where the streamlines are a
family of equal parabolas; this leads to an extension
of Blasius' result for uniform flow over a flat plate.
Finally, in the appendix, the skin friction coefficient
is calculated, from which is obtained a quantity anal_
ogous to the profile-drag coefficient of a swept wing,
for comparison with the results of Weber and Brebner
for the latter case. The comparison confirms the
assumptions made by them in the development of
their theory.





N-30898*

Royal Aircraft Establishment (Gt. Brit.)
NOTES ON THE RESPONSE OF A LINEAR VIBRA-
TION SYSTEM TO IMPACT LOADING. H. K. P.
Neubert. March 1954. 21p. diagrs. (RAE Tech.
Note Instn. 138)

In the course of investigations into the effect of im-
pact loadings, the general equations for deflection,
velocity, and acceleration of a linear system of one
degree of freedom with velocity damping have been
derived for a case of a mass being dropped onto a
spring (with negligible mass). Response curves are
plotted and discussed for two idealized cases of air-
craft landing, and for a drop table test. TIhe possible
errors for a proposed integrating touch-down velocity
meter are discussed. FIrequene,: spectra for three
typical impact curves have been computed and the
error introduced when measuring the impact with an
accelerometer of inadequate frequency response is
discussed.






MIISCELLANEOUS




NACA Rept. 1028

Errata No. 1 on "EFFECT OF ASPECT RATIO ON
THE AIR FORCES AND MOMENTS OF HARMON-
ICALLY OSCILLATING THIN RECTANGULAR
WINGS INJ SUPERSONIC POTENTIAL FLOW."
Charles E. Watkins. 1951.


NACA-Langley 6-22-54 4M


illUNIVERSliTY OF FLORIDA1

31262 08153 117 9


UNPUBLISHED PAPERS



N-1726f

ON SEVERAL MEASURES, TAKEN IN PRACTICE,
FOR ELIMINATING OSCILLATORY YAW WITH
FREE RUDDER. (Ujber einige praktisch
durchgeffihrte Massnahmen zur Beseitigung von
Gierschwingungen mit losem Ruder). R. Schmidt.
April 1954. 10p. diagrs., photo. (Trans. from
Lilienthal -Gesellschaft fur Luftfahrtforschung,
Berlin, Bericht 143, Nov. 6-7, 1941, p. 22-24)

Measures, tested and carried out on several air-
planes at Dornier works when unsatisfactorily
damped oscillatory yaw had occurred, and its elimi-
inationn by modifications of the tail had not been pos-
sible, are discussed. The two groups of measures
discussed are measures which augment the increase
of rudder moment in the region of zero position of
the controls; and measures which impede the free
motion relative to the air of the rudder.



N-31386*

MEASURES FOR ELIMINATION OF OSCILLATORY
YAW WITH FREE RUDDER. (Massnahmen zur
Beseitigung von Gierschwingungen mit losem
Seitenruder). H. G. Schumann. April 1954. 6p.
diagrs. (Trans. from Lilienthal-Gesellschaft fur
Luftfahrtforschung, Berlin, Bericht 143, Nov. 6-7,
1941, p. 27-29)

Experiences are given which were made with the
Ju 88 in controlling oscillatory yaw with free rudder.
This method of attaining the degree of rudder load
reduction required in the individual case by the con-
trol force itself, involved the return to harmless
variations of rudder-self-adjustment and eliminated
thus for a long time the difficulties occurring now
and again at other firms.



N-31449t

THEORETICAL AND EXPERIMENTAL STUDY OF
THE MEANS OF PROTECTING AN AIRPLANE
FROM GUST ACCELERATIONS. CHAPTER VIII -
EXPERIMENITAL STUDY. (Recherches the~oriques
et expkrimentales sur les moyens de soustraire un
avion aux accelbrations que peuvent engendrer les
perturbations atmosphiriques. Chapitre VIII -
Etude expi~rimentale). Rene Hirsch. February
1954. 45fp. diagrs., photos. (Trans. from
Ministere de l'Air. Publications Scientifiques et
Techniques, no. 138, 1938)

Changes in direction and intensity of the relative
wind which strikes the airplane in flight induce
changes in the total, hinge moment of the controlled
aileron so that the consecutive rotations of the latter
produce exactly defined variations of the total lift
coefficient. Two basic airfoils were studied, the
Potes 403 and NACA 23012.




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