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:
AA00009235:00054

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National Advisory Committee for Aeronautics



Research Abstracts
NO. 90 SEPTEMBER 27, 1955


CURRENT NACA REPORTS

NACA Rept. 1197

A STUDY OF THE CHARACTERISTICS OF HUMAN-
PILOT CONTROL RESPONSE TO SIMULATED AIR-
CRAFT LATERAL MOTIONS. Donald C. Cheatham.
1954. 11, 14p. diagrs., photos., tab. (NACA
Rept. 1197. Formerly RM L52C17)

There are presented studies of the characteristics of
pilot ability to control dynamically unstable yawing
oscillations, studies of pilot control response to
simulated aircraft yawing motions, and studies oi the
feasibility of representing pilot control response in
an analytical form.




NACA RM E55F28a /

STATISTICAL SURVEY OF 1 N fAURED
ON SCHEDULED AIRLINE F 1HTS VER THE / -
UNITED STATES AND CAN AIFOM NOVEMBEMt
1951 TO JUNE 1952. Porter ins.
September 1955. 44p. diagrs., oto
(NACA RM E55F28a) -

A statistical survey and a preliminary analysis are
made in an interim report of over 600 icing en-
counters obtained from a continuing program
sponsored by the NACA with the cooperation of the
airlines. Pressure-type icing-rate meters were in-
stalled on 11 airline aircraft of various types. Icing
conditions measured during scheduled operations
gave relative frequencies of liquid-water content,
icing rate, total ice accumulations, cloud tempera-
tures, as well as horizontal and vertical extent of
icing clouds. Liquid-water contents were higher
than data from earlier research flights in layer-type
coTds-butt-st~ghtty lower than previous data from
cumulus clouds.






NACA TM 1330

THEORY OF DYNAMIC CREEP. (K teorii
dinamicheskoi polzuchesti). A. A. Predvoditelev
and B. A. Smirnov. September 1955. 12p. diagr.
(NACA TM 1330. Trans. from Moscow Universitet,
Vestnik, v.8, no.8, 1953, p.79-86)

An analysis is given of the causes of the increase in
creep under varying loads. It is suggested that the
increase in creep is due to local rise in temperature


over the slip planes, thus facilitating slip. A theory
of dynamic creep is proposed based on theBecker -
theory of the after- flect, tdi8th treats the.rmetal as
a granular structure and ittnCdes a rate factor.
Comparison of the t eor th experit-entat results
is reserved for a fu ure apelr. ,., "


-L ..

NACA TN 3293

CUMULATIVE FATIGUE DAMAGE OF AXIALLY
LOADED ALCLAD 75S-T6 AND ALCLAD 24S-T3
ALUMINUM-ALLOY SHEET. Ira Smith, Darnley M.
Howard, and Frank C. Smith, National Bureau of
Standards. September 1955. 49p. diagrs., photos.,
5 tabs. (NACA TN 3293)

Results are presented of cumulative-fatigue-damage
tests made on 607 specimens machined from alclad
75S-T6 aluminum-alloy sheet 0.064 inch thick and
198 specimens of alclad 245-T3 and alclad 755-T6
aluminum-alloy sheet 0.032 inch thick. The tests of
the 0.064-inch-thick specimens 6lei'st ~5te 35 dif-
ferent loading conditions andln ,tl Ie.'4 32-
inch material consisted ofI'S3 d rent loa -;-n-
ditions.

.SEP s 19551

/
NACA TN 3294 ( 95

FRICTION STUDY OF AIRCRAFT"flREIAT IAL
ON CONCRETE. W. G. Hamrite, Boeing Arfplane
Company. September 1955. 34p. diagrs., photos.
(NACA TN 3294)

A systematic study was made of the variation of
frictional resistance between typical tire-tread
material and three concrete surfaces of different
roughness at various temperatures and normal
pressures. The tire-tread specimens were taken
from the thickest portion of worn ten-ply tires, and
the three concrete test blocks were poured from the
same mix but subjected to dilerent surface finishes.
Curves are presented ol the apparent coefficient of
friction as a function of normal pressure.




NACA TN 3477

HYDRODYNAMIC PRESSURE DISTRIBUTIONS OB-
TAINED DURING A PLANING INVESTIGATION OF
FIVE RELATED PRISMATIC SURFACES. Walter J.
Kapryan and George M. Boyd, Jr. September 1955.
82p. diagrs., photos., 5 tabs. (NACA TN 3477)


*AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1519 H ST., NW., WASHINGTON 25, D C., CITING CODE NUMBER ABOVE EACH TITLE.
THE REPORT TITLE AND AUTHOR.


:t? "/3 36
p' c-eP w







2




Hydrodynamic pressure distributions have been ob-
"tained during pure planing for five related prismatic
surfaces. The distributions gave integrated lifts
that in almost every case were well within 10 percent
of the applied load. Comparison of experiment with
theory shows that existing theories will adequately
predict flat-plate pressures. For the V-shaped sur-
faces, experiment and theory are in poor agreement.
The lift and center-of-pressure data for both the flat
and V-shaped surfaces are in good agreement with
recent experimental and theoretical NACA research
on planing surfaces.


NACA TN 3479

ANALYSIS OF THE HORIZONTAL-TAIL LOADS
MEASURED IN FLIGHT ON A MULTIENGINE JET
BOMBER. William S. Aiken, Jr. and Bernard
Wiener. September 1955. i, 69p. diagrs., photo.,
6 tabs. (NACA TN 3479)

Horizontal-tail loads were measured in gradual and
abrupt longitudinal maneuvers on two configurations
of a four-engine jet bomber. The results obtained
have been analyzed to determine the flight values
of the coefficients important in calculations of hori-
zontal tail loads. The least-squares procedure used
to determine aerodynamic tail loads from strain-
gage measurements of structural tail loads which
were affected by temperature is covered in detail.
The effect of fuselage flexibility on the airplane
motion is considered in the analysis of the abrupt-
maneuver data. When possible, wind-tunnel results
are compared with flight results. Some calculations
of critical horizontal-tail loads beyond the range of
the tests are given and compared with design loads.



NACA TN 3486

MEASUREMENTS OF TURBULENT SKIN FRICTION
ON A FLAT PLATE AT TRANSONIC SPEEDS.
Raimo J(aakko) Hakkinen, California Institute of
Technology. September 1955. 41p. diagrs., photo,
tabs. (NACA TN 3486)

The design and construction of a floating-element
skin-friction balance are described. This instru-
ment was applied to measurements of local skin
friction in the turbulent boundary layer of a smooth
flat plate at high-subsonic Mach numbers and super-
sonic Mach numbers up to 1. 75. The principal
difficulties which exist in comparing skin-friction
coefficients at various Mach numbers are discussed.



NACA TN 3491

EXPERIMENTAL INVESTIGATION OF ECCENTRI-
CITY RATIO, FRICTION, AND OIL FLOW OF LONG
AND SHORT JOURNAL BEARINGS WITH LOAD-
NUMBER CHARTS. G(eorge) B. DuBois, F(red) W.
Ocvirk, and R. L. Wehe, Cornell University.
September 1955. 63p. diagrs., tabs. (NACA TN
3491)


NACA
RESEARCH ABSTRACTS NO.90



The performance of plain bearings under steady
central loading are compared and sunniarized by
single-line curves covering the range of length-
diameter ratios both abore and below I. Experi -
mental date on eccentricity ratio, friction, and oil
flow for length-diameter ratios of 1, 1-1 2, and 2
are shown for comparison .'ith earlier data for
length-diameter ratios of 1.4, 1 2. and 1. The
combined data provide charts of plain-oearing per-
formance which cover the practical range of length-
diameter ratio.



NACA TN 3493

DEVELOPMENT OF EQUIPMENT AND OF EXPERI-
MENTAL TECHNIQUES FOR COLUMN CREEP
TESTS. Sharad A. Patel, Martin Bloom, Burton
Erickson, Alexander Chaick and N(icholas) J(ohn)
Hoff, Polytechnic Institute of Brooklyn. September
1955. 20p. diagrs.. photos., tab. INACA TN 3493)

Equipment and procedures developed for testing
aluminum-alloy columns subjected to constant loads
at elevated temperatures are described. Particular
emphasis was put on determination of the influence
of initial dei nations from straightness on the critical
time of the column, that is, the time necessary for
the column to buckle %hen subjected to a constant
load. Results are presented of tests of a number of
2024-T4 aluminumn-alloy columns having large slen-
derness ratios.



NACA TN 3503

REDUCTION OF PROFILE DRAG AT SUPERSONIC
VELOCITIES BY THE USE OF AIRFOIL SECTIONS
HAVING A BLUNT TRAILING EDGE. Dean R.
Chapman. September 1955. 29p. diagrs., photo.
(NACA TN 3503. Supersedes RM A9H11)

A preliminary theoretical and experimental investi-
gation has been made on the aerodynamic character-
istics of blunt-trailing-edge airfoils at supersonic
velocities. The theoretical considerations indicate
that properly designed airfoils with moderately blunt
trailing edges can have less profile drag, greater
lift-curve slope, and a higher maximum lift-drag
ratio than conventional sections. These predictions
have been substantiated by experimental measure-
ments on airfoils of 10-percent-thickness ratio at
Mach numbers of 1.5 and 2.0, and at Reynolds num-
bers between 0. 2 and 1. 2 million.



NACA TN 3514

RESPONSE OF HOMOGENEOUS AND TWO-
MATERIAL LAMINATED CYLINDERS TO SINUSOI-
DAL ENVIRONMENTAL TEMPERATURE CHANGE,
WITH APPLICATIONS TO HOT-WIRE ANEMOM-
ETRY AND THERMOCOUPLE PYROMETRY.
Herman H. Lowell and Norman (A.) Patton.
September 1955. ii, 143p. diagrs., tabs. (NACA
TN 3514)







NACA
RESEARCH ABSTRACTS NO. 90



A theoretical investigating ot the response ot homno-
geneous and t.o-niaterial laminated, iiiiiiite cylin-
ders to sinusoidal environnenrtal temperature and
or small heat-transfer coefficient changes was made.
Generalized results are given ior the cylinder con-
sisting ol a shell of high thermal conductivity and a
core of Io., conduct i. it.' The Othatnor ol a nuliber
of specific platinurm-fused-quartz ires" no '.arirng
construction and diameter exposed to a representa-
tive airstream is indicated. For ratios l metal
tnickness to over-all radius of 0. 1, response ampli-
tude gains of aDout 4. 5 are predicted as compared
with gains of more than 10 for infinitesimal shells.
For a relative shell thickness of. 05.0 frequency re-
sponses at hot-.ire aneiior..eters, exposed-wire re-
sistance thermometers, or thermocouples Aould be
extended by at least an order ol magnitude. Simnpli-
fied analyses are included which are not exact but
are adequate lor design use.


NACA TN 3522

MEASUREMENTS OF THE EFFECTS OF FINITE
SPAN ON THE PRESSURE DISTRIBUTION OVER
DOUBLE-WEDGE WINGS AT MACH NUMBERS
NEAR SHOCK ATTACHMENT. Walter G. Vincenti.
September 1955. 50p. diagrs. (NACA TN 3522)

Results are presented of measurements at low super-
sonic speeds of the pressure distribution on t1o
wings having a common double-wedge section and
aspect ratios 2 and 4. Comparable results for as-
pect ratio infinity have been published in NACA TN
3225. The results cover the Mach number range
from 1.166 to 1.377, which brackets the value (1.221)
for bow-wave attachment at zero angle of attack.
The data are discussed and compared with the previ-
ous two-dimensional findings.



NACA TN 3523

THE EFFECTIVENESS OF WING VORTEX GENERA-
TORS IN IMPROVING THE MANEUVERING CHARAC-
TERISTICS OF A SWEPT-WING AIRPLANE AT
TRANSONIC SPEEDS. Norman M. McFadden,
George A. Rathert, Jr., and Richard S. Bray.
September 1955. 43p. diagrs., photos., tab.
(NACA TN 3523. Supersedes RM A51J18)

The effects of wing vortex generators, multiple
boundary-layer fences, and extension of the outer
two segments of the wing leading-edge slats on the
aerodynamic characteristics ot a 350 swept-wing
fighter were measured in (light tests at transonic
speeds and high altitudes. Significant improvements
were obtained in the pitch-up and wing-dropping-
tendency characteristics with certain arrnagements
of vortex generators.



NACA TN 3562

VARIATION OF BOUNDARY-LAYER TRANSITION
WITH HEAT TRANSFER ON TWO BODIES OF
REVOLUTION AT A MACH NUMBER OF 3.12.
John R. Jack and N. S. Diaconis. September 1955.
16p. diagrs., photos. (NACA TN 3562)


Cooling a cone-cylinder model to a wall-to-free-
stream ratio of approximately 1.4 increased the
transition Reynolds number from a value of 2.0 x 106
at equilibrium to 10.6 x 106. For temperature
ratios less than 1.4, the boundary-layer flow was en-
tirely laminar. For a parabolic-nosed body, the
transition Reynolds number was about twice that of
the cone-cylinder model over the temperature range
investigated.




NACA TN 3563

HEAT LOSS FROM YAWED HOT WIRES AT SUB-
SONIC MACH NUMBERS. Virgil A. Sanaborn and
James C. Laurence. September 1955. 44p.
diagrs., photo. INACA TN 3563)

Heat-loss data at angles of yavw and fixed subsonic
Mach numbers for several vires of different diam-
eters commonly used in hot-mire anenomnetry are
presented. Possible methods of correlating the
data are examined. The relation oi the Reynolds
number normal to the flo.., which has been used by
most researchers, was inadequate except near a
Mach number of zero. An empirical relation based
on weighted addition of the heat losses of wires
normal and parallel to the flow correlated all data
reasonably well.



NACA TN 3566

A POLAR-COORDINATE SURVEY METHOD FOR
DETERMINING JET-ENGINE COMBUSTION-
CHAMBER PERFORMANCE. Robert Friedman and
Edward R. Carlson. September 1955. 29p.
diagrs., photo., tab. (NACA TN 3566)

An automatic polar-coordinate traversing system is
described that sweeps a probe through a quarter-
annular exhaust duct circumferentially at selected
radial positions. With a single combined pressure
and temperature probe, temperature and pressure
are recorded simultaneously as a function of probe
position. The use of these data in calculating
temperature and flow profiles, combustion efficiency,
and pressure loss is shown.








BRITISH REPORTS






N-38605*

Aeronautical Research Council (Gt. Brit.)
THE USE OF QUARTZ IN THE MANUFACTURE OF
SMALL DIAMETER PITOT TUBES. J. R. Cooke.
1955. 14p. diagrs., photos., tab. (ARC CP 193)











This note describes the method of manufacture of
small quartz-tipped pitot tubes (down to 0.005 in.
outside tip diameter) which have been used success-
fully for boundary-layer measurements on small
models in a supersonic wind tunnel. Tests have
been made of the effects of taper and end finish on
the accuracy of measurement, and of the effect of
the inside diameter of the tip (for a standard taper)
on response rate. For a given inside tip diameter
the tapered quartz tubes gave a faster response rate
than the stainless steel hypodermic tubes previously
used.




N-38606*

Aeronautical Research Council (Gt. Brit.)
A NOTE ON THE SOUND FROM WEAK DISTURB-
ANCES OF A NORMAL SHOCK WAVE. Alan PowelL
1955. 10p. diagrs. (ARC CP 194)

The disturbances of a shock wave by sound waves or
temperature fluctuations are studied in one dimen-
sion to a first-order approximation. In general,
both sound waves and temperature fluctuations arise
behind the shock wave. Expressions are given for
their amplitudes and calculated for y = 1.4. Sound
waves colliding with the shock wave are amplified,
but sound waves are almost annihilated by weak
shock waves if originally travelling in the same di-
rection as the shock wave. Small temperature
fluctuations give rise to much sound on an acoustical
scale.




N-38607*

Aeronautical Research Council (Gt. Brit.)
REQUIREMENTS FOR UNIFORMITY OF FLOW IN
SUPERSONIC WIND TUNNELS. D. E. Morris and
K. G. Winter. 1955. 9p. diagr. (ARC CP 197)

An analysis is made of the effects of nonuniformity
of flow on the pressure measurements on the surface
of a model and also on the force and moment meas-
urements. The following standards of flow uniform-
ity are derived variations in flow direction to be
less than 0.10 in the range M = 1.4 to 3; variation
in Mach number to be less than 0.003 at M = 1.4
increasing to 0.01 at M = 3. A brief analysis is
made of the errors in model manufacture and their
effects on force and pressure measurements. Using
the same standards as were used in deducing the
requirements for flow uniformity quoted above, it is
concluded that present standards of model manu-
facture are satisfactory overall, though for accurate
pressure plotting tests at low supersonic Mach num-
bers a higher standard is desirable.



N-38608*

Aeronautical Research Council (Gt. Brit.)
A CRITERION FOR THE PREDICTION OF THE RE-
COVERY CHARACTERISTICS OF SPINNING AIR-
CRAFT. T. H. Kerr. 1955. 22p. diagrs., tabs.
(ARC CP 195)


NACA
RESEARCH ABSTRACTS NO. 90



It has been deduced that the t. o miost in.portant pa-
rameters are the unbalanced rolling-nion.ent coeffi-
cient about the wind axis in the spin and the ratio of
pitching to rolling moment of inertia. Using the
results of full-scale spinning tests on 33 aircraft, it
has been possible to establish empirical relation-
ships between the estimated unbalanced rolling-
moment coefficient and the inertia ratio which effect-
ively divide the aircraft into the three groups which
have satisfactory, borderline, and unsatisfactory re-
covery characteristics. A simple method is pre-
sented for estimating the unbalanced rolling-moment
coefficient knowing only the shape of the aircraft.
The empirical relationships should give a good indi-
cation of the spin-recovery characteristics on new
designs.

N-38616*

Aeronautical Research Council (Gt. Brit.)
MODEL TESTS ON THE EFFECTS OF SLIPSTREAM
ON THE FLOW AT VARIOUS TAILPLANE POSI -
TIONS ON A FOUR-ENGINED AIRCRAFT. PART I.
TESTS WITH CONTRA-ROTATING PROPELLERS.
D. E. Hartley, A. Spence, and D. A. Kirby.
PART II. TESTS WITH SINGLE ROTATING PRO-
PELLERS. D. A. Kirby. 1955. 37p. diagrs.,
tabs. (ARC R & M 2747; ARC 12, 355; ARC 14, 166.
Supersedes RAE Aero 2322; RAE Aero 2322a)

Systematic wind-tunnel tests have been made to in-
vestigate the effects of slipstream on the flow near
the tail plane of a typical civil transport with four
contra-rotating propellers. Tall-plane height has
been varied for each of several wing-body arrange-
ments; only one tail plane and one propeller position
have been used. This report presents the main re-
sults in the form of changes in mean downwash angle
and velocity at the tail plane, as functions of tail-
plane position, lift coefficient, and propeller thrust.


N-38617*

Aeronautical Research Council (Gt. Brit.)
DETERMINATION OF THE STRESS DISTRIBUTION
IN REINFORCED MONOCOQUE STRUCTURES.
PART I. A THEORY OF FLAT-SIDED STRUC-
TURES. L.S.D. Morley. 1955. 23p. dagrs.,
photos. (ARC R &M 2879; ARC 14,814. Superse-
des RAE Structures 120)

This paper is concerned with the estimation of the
stress distribution in the neighborhood of a discon-
tinuity in reinforced monocoque flat-sided struc-
tures. A theory is given based upon a shell model
possessing uniformly distributed stringers but dis-
crete ribs, which can serve as a basis for the prac-
tical solution of a wide range of flat-sided struc-
tures such as rectangular or polygonal fuselages and
wing boxes.


N-38618*

Aeronautical Research Council (Gr. Brit.)
THE THEORETICAL WAVE DRAG OF SOME
BODIES OF REVOLUTION. L. E. Fraenkel. 1955.
26p. diagrs., tab. (ARC R & M 2842; ARC 14, 334.
Supersedes RAE Aero 2420)







NACA
RESEARCH ABSTRACTS NO. 90



This report investigates the wave drag of bodies of
revolution with pointed or open-nose forebodies and
pointed or truncated afterbodies. The "quasi-
cylinder" and"slender-body" theories are reviewed,
a reversibility theorem is established, and the con-
cept of the interference effect of a forebody on an
afterbody is introduced. The theories are applied
to bodies whose profiles are either straight or para-
bolic arcs, formulas and curves being given for
forebody and afterbody drag, and for the interfer-
ence drag. The results of the tAo theories are com-
pared and are seen to agree well in the region of
geometries where both theories are applicable.





N-38619*

Aeronautical Research Council (Gt. Brit.)
AN EXPERIMENTAL INVESTIGATION OF STRESS
DIFFUSION IN NON-BUCKLING PLATES. L. H.
Mitchell. 1955. 20p. diagrs., photos. (ARC
R & M 2878. Supersedes ARC 14,934; Strut 1540)

This report provides experimental results for com-
parison with theoretical analyses of stress diffusion
problems. The structures considered consist of
plane reinforced sheet which has been assumed not
to buckle. Symmetrical loads are applied to the
edge booms connected to the sheet by continuous no-
slip joints. Attention is concentrated on the stress
distribution near the ends of the parallel strips of
plate. An outline of the existing theoretical work
which' is applicable to this type of problem is given.
The stringer-sheet theory is compared with the
photoelastic results. Some attention is also given
to transverse end stiffeners which seem to have
little effect on the shear stresses.






N-38620*

Aeronautical Research Council (Gt. Brit.)
THE BOUNDARY LAYER WITH DISTRIBUTED SUC-
TION. M. R. Head. 1955. 100p. diagrs., photos,
tabs. (ARC R & M 2783. Supersedes ARC 13, 897;
FM 1547; Perf. 771)

Experiments performed in flight at Reynolds num-
bers in the region of 3 x 106 have clearly demon-
strated the stabilizing effect of small amounts of
distributed suction on the laminar boundary layer.
In the absence of a pressure gradient and in adverse
gradients similar to those occurring on a normal
airfoil, transition of the boundary layer to the tur-
bulent form has been prevented by the use of such
suction quantities as may be expected to lead to very
considerable reductions in effective drag. It ap-
pears, however, that for extensive laminar flow to
be achieved in this way, the surface must be free
from such excrescences as would cause transition
in the absence of suction.


N-38621*

Aeronautical Research Council (Gt. Brit.)
METHODS FOR CALCULATING THE LIFT DISTRI-
BUTION OF WINGS (SUBSONIC LIFTING-SURFACE
THEORY). H. Multhopp. 1955. 96p. diagrs..
tabs. (ARC R & M 2884; ARC 13,439. Supersedes
RAE Aero 2353)
These methods for calculating the load distribution
on wings of any plan forn are based on the concep-
tions of lifting-surface theory. Computer work
time is shortened by careful choice of the positions
of pivotal points, by plotting once for all those parts
of the doAnwash integral which occur frequently and
by a consequent application of approximate integra-
tion methods similar to those devised by the author
for lifting-line problems. The basis of the method
is to calculate the local lilt and pitching moment at
a number of chordwise sections from a set of linear
equations satisfying the donnvash conditions at two
pivotal points in each section.





N-3871

Aeronautical Research Council (Gt. Brit.)
SIMPLE EVALUATION OF THE THEORETICAL
LIFT SLOPE AND AERODYNAMIC CENTRE OF
SYMMETRICAL AEROFOILS. H. C. Garner.
1955. 20p. tabs. (ARC R & M 2847. Supersedes
ARC 14,337; Perf.847; S & C 2561)

This paper presents a simple method of calculating
theoretical values of the lift slope (al)T and the
position of aerodynamic center hT in two-
dimensional incompressible flow. Starting with the
ordinates of an airfoil, the method in section 3 pro-
vides first and second approximations to both de-
rivatives, which are compared with exact theory
and other calculated values in Tables 2 and 3 for
various symmetrical airfoils listed in Table I. In
section 5 a correction to the first approximation is
introduced so as to permit the evaluation of (al)T
within 1,2 percent and hT within about 0.001 in
less than a quarter of an hour. A complete illus-
trative calculation is set out in Table 4.





N-38712'

Aeronautical Research Council (Gt. Brit.)
BOUNDARY-LAYER CONTROL FOR HIGH LIFT BY
SUCTION AT THE LEADING-EDGE OF A 40 DEG
SWEPT-BACK WING. E. D. Poppleton. 1955.
38p. diagrs., tabs. (ARC R & M 2897; ARC 14,771.
Supersedes RAE Aero 2440)

Wind-tunnel tests on the 10-percent-thick, constant-
chord, aspect-ratio-4.6 wing are discussed.
Boundary-layer control was applied along the whole
leading edge; a comparison was made between the
effects of distributed suction and suction through a










slot. A 45-percent Fowler flap was used in some
tests. The overall effect of the two systems was
similar, giving an increase in CLmax by increas-
ing the stalling angle of attack and making the wing
statically stable up to the stall, when there was a
severe loss of lift. The tests were designed to de-
termine whether leading-edge suction would produce
comparable increases in CLmax on swept wings
and, also whether tip stall could be prevented.





N-38713*

Aeronautical Research Council (Gt. Brit.)
ON THE APPLICATION OF OBLIQUE CO-
ORDINATES TO PROBLEMS OF PLANE ELAS-
TICITY AND SWEPT-BACK WING STRUCTURES.
W. S. Hemp. WITH AN APPENDIX. S. R. Lewis.
1955. 46p. diagrs., tabs. (ARC R & M 2754; ARC
12,981. Supersedes College of Aeronautics Rept.
31; College of Aeronautics Rept. 44)

Methods are discussed by which designers can solve
problems of stress distribution and deflection for
the case of sweptback wing structures whose ribs
lie parallel to the direction of flight. The mathe-
matical basis is developed and formulas are derived.
The results are applied to a uniform, symmetrical,
rectangular section sweptback box. Theories of
stress distribution and deflections are obtained for
the case of loading by normal forces and couples
applied to the ends of the box. The main results
are then generalized to cover the case of a more
representative wing structure. Functions useful in
the application of the theory are given in an appendix.






N-38714*

Aeronautical Research Council (Gt. Brit.)
LOW-SPEED TUNNEL MODEL TESTS ON TAIL-
PLANE ROLLING MOMENTS IN SIDESLIP.
A Spence, J. W. Leathers, and D. A. Kirby. 1955.
20p. diagrs., tabs. (ARC R& M 2941; ARC 14,701.
Supersedes RAE Tech. Note Aero 2123)

Measurements were made of the effect of sideslip on
the rolling moment on a 41.50 sweptback tail plane
mounted at three heights on the fin of a model of a
single jet aircraft with a 400 sweptback wing. Inci-
dence and tail-plane setting were varied, and the ef-
fects of rudder deflection were obtained with the
tail plane at the top of the fin. Brief results on a
delta aircraft model with a delta tail plane at the top
of the fin are also included. Values of the rolling
moment on the tail plane were obtained from meas-
urements of the bending moment on the starboard
half of the tail plane about a hinge just outside the
fin.


NACA
RESEARCH ABSTRACTS NO. 90



N-38715*

Aeronautical Research Counc II IGt. Brit.)
TWO-DIMENSIONAL CONTROL CHARACTER-
ISTICS. L. W. Bryant, A. S. Halliday, and A. S.
Batson. 1955. 47p. diagrA. IARC R M 2730.
Supersedes ARC 13,039; S & C 2385: ARC 13.065;
S & C 2386)

Researches on the lift, pitching moments, and hinge
moments of airfoils with plain flaps have been car-
ried out at the National Physical Laooratory at a
Reynolds number of about 106. The results have
been presented in a generalized form, *hich shows
promise of being applicable over a wide field. It
appears that a suggestion due to Preston that the
ratio of experimental lift slope IdCL dy = al) to the
theoretical value (al)T, corresponding to the
Joukowsky condition of flo* past the trailing edge,
provides a criterion giving the combined effects of
Reynolds number, transition points, and airfoil
shape on dCL/dct, and is a ivry useful starting
point for the estimation of control characteristics.





N-38716*

Aeronautical Research Council IGt. Brit.)
PERMISSIBLE DESIGN VALUES AND VARIABILITY
TEST FACTORS. R. J. Atkinson. 1955. 20p.
diagrs., tabs. (ARC R & M 2877; ARC 11,619;
ARC 13,748. Supersedes RAE Tech. Note Structures
15; RAE Tech. Note Structures 61)

For the design of structural elements it is postulated
that: not more than 10 percent of any given design
should have strength below the design value, and not
more than 0.1 percent should have strength below 90
percent of the design value. This rule forms a
working basis for the interpretation of tests on sta-
tistical lines. On the basis of a fixed probability the
report deduces: expressions for the derivation of
permissible design values from a given number of
test results, the number of test results required so
that the estimates of permissible design values can
be regarded as sufficiently accurate, and the factor
which should be applied to the results of tests on any
number of similar components designed to meet a
specified requirement.






N-38717*

Aeronautical Research Council iGd. Brit.)
IMPROVEMENTS IN THE FATIGUE STRENGTH OF
JOINTS BY THE USE OF INTERFERENCE FITS.
W. A. P. Fisher and W. J. Winworth. 1955. 17p.
diagrs., photos., tabs. IARC R & M 2874: ARC
15,014. Supersedes RAE Structures 127)








NACA
RESEARCH ABSTRACTS NO. 90



Fatigue test results are given for aluminum alloy
flat bars with a single hole loaded by a pin in double
shear. In one series the pin was fitted directly in
S the hole with various degrees of interference fit up
to 0.003 in. excess diameter. The other series had
a mild steel bush interposed sith similar degrees of
interference in the bar, but with a push fit between
pin and bush. Both sets showed a great increase in
fatigue strength for interference fits above a critical
value.






N-38718*

Aeronautical Research Council (Gt. Brit.)
AN EXAMINATION OF THE FLOW AND PRESSURE
LOSSES IN BLADE ROWS OF AXIAL-FLOW TUR-
BINES. D. G. Ainley and G. C. R. Mathieson.
1955. 33p. diagrs. (ARC R & M 2891; ARC
14,232. Supersedes NGTE R.86)

Available information is studied and analyzed to de-
termine magnitudes of gas pressure losses and de-
Ilections in a wide variety of blade rows and to de-
termine the separate influences of variables such as
blade shape, blade spacing, gas Mach number,
Reynolds number, incidence, etc. Special attention
is paid to "secondary losses. Effects of blade tip
clearance are also considered. Empirical guiding
rules and charts are derived from which approximate
values of the overall pressure losses and gas deflec-
tions in a range of blade rows can be deduced. It is
found that secondary losses can in many instances be
large, the loss being generally found to be great
when the blading has low reaction.






N-38719*

Aeronautical Research Council (Gt. Brit.)
FLUTTER AND RESONANCE CHARACTERISTICS
OF A MODEL CANTILEVER WING CARRYING
LOCALISED MASSES. N. C. Lambourne. 1955.
25p. diagrs., tabs. (ARC R & M 2866. Supersedes:
ARC 13,910; 0.939; ARC 11,008; 0.687)

Resonance tests on a model cantilever wing carrying
concentrated masses were made in conjunction with
flutter tests. Measurements were made with
masses up to approximately five times the mass of
the bare wing added at two positions. Flutter and
resonance characteristics are placed in juxtaposi-
tion. An attempt is made to correlate the two sets
of phenomena by means of the Kiissner criterion.
Distortion modes of flutter are analyzed into normal
mode components. Results suggest that for a wing
rigidly fixed at the root and carrying a single con-
centrated mass the first three normal modes are
sufficient to define the flutter mode.
Copies obtainable from NACA, Washington


7




N-38720'

Aeronautical Research Council (Gt. Brit.)
SOME APPLICATIONS OF THE LAME FUNCTION
SOLUTIONS OF THE LINEARISED SUPERSONIC
FLOW EQUATIONS. PART I FINITE SWEPT-
BACK WINGS WITH SYMMETRICAL SECTIONS AND
ROUNDED LEADING EDGES. PART I CAMBER-
ED AND TWISTED WINGS. G. M. Roper. 1955.
42p. diagrs. (ARC R & M 2865; ARC 14,473; ARC
14,475; ARC 14,476. Supersedes RAE Aero 2436;
RAE Aero 2437)

In the present paper some special solutions are
found. Some of these solutions are combined with
previous solutions to give (a) pressure distribution
and wave drag at zero lift on some finite unyawed
sweptlack ings having symmetrical sections with
rounded leading edges and wing tips perpendicular to
the wind direction, and (b) the change in pressure
distribution and wave drag at zero lift on the surface
of a Sqwure wing when the thickness chord ratio is
modiiied. Some additional solutions applicable to
cambered and twisted wings are also given.







N-38721'

Aeronautical Research Council (Gt. Brit.)
THE APPLICATION OF THE EXACT METHOD OF
AEROFOIL DESIGN. M. B. Glauert. 1955. 45p.
diagrs., tabs. (ARC R & M 2683. Supersedes
ARC 10,933; FM 1161)

This report considers in detail the design of air-
foils by Lighthill's exact method, in which the ve-
locity over the airfoil surface is prescribed as a
function of the angular coordinate on the circle into
which the airfoil may be transformed. The mathe-
matical basis of the method is set out, means for
obtaining desired characteristics for the airfoil are
developed, and the procedure to be followed in the
actual design is fully discussed. Various special
functions are introduced to increase the range and
practical utility of the velocity distributions obtain-
able, and these and other functions are fully tabu-
lated. The calculations for the design of a particu-
lar thick suction airfoil are set out in detail.
Copies obtainable from NACA, Washington






N-38722*

Aeronautical Research Council (Gt. Brit.)
AN EXPERIMENTAL INVESTIGATION OF THE
BOUNDARY LAYER ON A POROUS CIRCULAR
CYLINDER. D. G. Hurley and B(rian) Thwaites.
1955. 14p. diagrs., photos. (ARC R & M 2829.
Supersedes ARC 14,158; FM 1584)











The report describes an experimental investigation
of the boundary layer on the surface of a porous
circular cylinder at which there is a normal inward
velocity. The primary object of the experiments
was to test the approximate theory of reference 1 for
calculating the development of a laminar boundary
layer under conditions of continuous suction. The
formula given in that reference for calculating the
momentum thickness of the layer gave results in ac-
cord with the experimental determinations. Owing
to practical difficulties in the exploration of the very
thin boundary layers and in the determination of the
velocity gradient around the surface, other com-
parisons with the theory were difficult.




N-38723*

Aeronautical Research Council (Gt. Brit.)
FORMULAE FOR ESTIMATING THE FORCES IN
SEAPLANE-WATER IMPACTS WITHOUT ROTA-
TION OR CHINE IMMERSION. R. J. Monaghan and
P. R. Crewe. 1955. 28p. diagrs., tabs. (ARC
R & M 2804; ARC 12,399. Supersedes RAE
Aero 2308)

This report contains design formulas for estimating
the maximum forces, together with the times and
drafts associated with these forces, in main-step
landings of seaplanes provided there is neither ro-
tation nor chine immersion. Good agreement is
formed with the results of model tests made under
controlled conditions at NACA. The basic formulas
and curves presented are considered to be the most
satisfactory and accurate of the many proposed in
recent years. They involve the use of a new basic
parameter which is a measure of the effect of
forward velocity; a new formula for associated
mass, and a new method of plotting which is con-
sidered to be the most useful for the analysis of ex-
perimental data.




N-38724*

Aeronautical Research Council (Gt. Brit.)
WIND-TUNNEL TESTS ON THE NACA 63A009
AEROFOIL WITH DISTRIBUTED SUCTION OVER
THE NOSE. N. Gregory and W. S. Walker. 1955.
17p. diagrs., tabs. (ARC R & M 2900. Supersedes
ARC 15,184; Perf. 987; FM 1787)

The effects of distributed suction on the stalling
characteristics of the airfoil are described. The
most economical extent of suction was from the lead-
ing edge for 2.75 percent chord round the upper
surface. At a R = 1.15 x 106, a suction-quantity
coefficient of 0.0034 increased CLmax from 0.86
to 1.50 by delaying the stall from a = 110 to
a = 200. Scale effect on the flow was investigated
at a = 14. The airfoil was also tested with a
20-percent split flap at 600 deflection. Suction
gave half the increase on the flapped airfoil that it
gave on the plain airfoil. The airfoil was modified
for further testing by reducing the chord and
blunting the nose.


NACA
RESEARCH ABSTRACTS NO. 90


N-38725*

Aeronautical Research Council (Gl. Brit.)
DETAILED OBSERVATIONS MADE AT HIGH IN-
CIDENCES AND AT HIGH-SUBSONIC MACH NUM-
BERS ON GOLDSTEIN 1442/1547 AEROFOIL.
H. H. Pearcey and M. E. Faber. 1954. 52p.
diagrs., photos., tabs. (ARC R & M 2849. Super-
sedes ARC 13,531; FM 1498; Perf. 714)

Surface-pressure distribution, shock-wave photo-
graphs, and observations of boundary-layer separa-
tion have been made over a wide range of angle of
attack. The observations enable the effects of
compressibility on CLmax and on the nature of the
stall to be studied in detail for the two-dimensional
case. The pitching-moment coefficients, also, can
be integrated from the pressure distributions. Cer-
tain features of the results are thought to be of fairly
general interest and application.







N-38728*

Royal Aircraft Establishment (Gt. Brit.)
TECHNIQUES FOR THE MEASUREMENT OF THE
AERODYNAMIC FORCES ON OSCILLATING AERO-
FOILS. W. G. Molyneux. June 1955. 30p. diagrs.
(RAE Tech. Note Structures 161)

The various techniques for oscillatory force meas-
urements are considered in relation to their applica-
tion to the measurement of the aerodynamic coeffi-
cients for a rectangular wing oscillating in modes of
vertical translation and uniform pitch. It is shown
that the eight relevant coefficients Lz, Lz, L,,
Li, Mz, M., My and M& are obtainable by any
of the techniques described. The survey is not ex-
haustive, but it provides a basis for comparison of
the various techniques and should be of assistance to
investigators in this field in indicating the particular
technique most likely to meet their requirements.






N-38729*

Royal Aircraft Establishment (GL. Brit.)
THE EFFECT OF WATER ON THE POROSITY OF
PARACHUTE FABRICS. J. E. Swallow. May
1955. 18p. diagrs., tabs. (RAE Tech. Note
Chem. 1248)

Air flow through parachute fabrics was found to be
seriously affected by water. The porosity of the
nylon, cotton, Fortisan and Terylene fabrics ex-
amined was decreased and became negligible for the
closer weaves. This was mainly a surface tension
effect, but swelling was a contributory factor for
cellulosic fabrics. Mock-leno weave nylon fabrics
were least affected.







NACA
RESEARCH ABSTRACTS NO. 90


N-387301

Royal Aircraft Establishment (Gt. Brit.)
ON THE INTEGRAL EQUATIONS OF TWO DIMEN-
SIONAL SUBSONIC FLUTTER DERIVATIVE
THEORY. D. E. Williams. June 1955. 39p.
(RAE Structures 181)

This note gives the result of an attempt to find an
analytical solution of Possio's integral equation -
the equation which connects the downwash and the
pressure distribution on an airfoil oscillating in two-
dimensional subsonic conipressible flow. A method
is given for solving this problem and for solving the
corresponding problem in incompressible flow the
solution of Birnbaum's integral equation.




N-38732*

Royal Aircraft Establishment (Gt. Brit.)
THE DETERMINATION OF FLUORINE IN ORGANIC
COMPOUNDS CONTAINING FLUORINE AND PHOS-
PHORUS. T. R. F. W. Fennell. May 1955. lip.
diagr., tabs. (RAE Tech. Note Chem. 1251)

A published method for the determination of fluoride
in the presence of phosphate ion has been found to
yield erroneous results. The method has been
modified to overcome this fault.




N-38759*

Aeroplane and Armament Experimental Establish-
ment (Gt. Brit.) THE EFFECT OF THE GROUND
ON A HELICOPTER ROTOR IN FORWARD FLIGHT.
I. C. Cheeseman and W. E. Bennett. July 11, 1955.
13p. diagrs. (AAEE'Res/288)

An approximate method of estimating the effect of
the ground on the lift of a rotor at any forward speed
is described. Flight tests on several different air-
craft show reasonable agreement with the theory.
Curves are given showing the relation between
thrust, height, speed, and power. The theory has
been extended to include the effect of a variation in
blade loading and shows that within the range that
this parameter takes on present single rotor heli-
copters the effect is small.




N-38761

Royal Aircraft Establishment (Gt. Brit.)
A UNIFIED THEORY OF PERFECTLY PLASTIC
PLATES. E. H. Mansfield. May 1955. 53p.
diagrs. (RAE Structures 170)

A theory is developed for determining the collapse
load and the collapse mechanism for perfectly plas-
tic plates under normal loading. A number of solu-
tions to simple problems is first presented and the
theory is extended to deal with plates of arbitrary
plan carrying a concentrated load, and to plates of
rectangular or regular polygonal plan carrying a
uniformly distributed load.


N-38781

Forest Products Research Lao. (Gt. Brit.)
INVESTIGATIONS INTO GLUES AND GLUING.
PROGRESS REPORT EIGHTY-FIVE JUNE 1955.
BEHAVIOR OF GLUED WOOD PRODUCTS IN LIGHT
NAVAL CRAFT. PART I SYNOPTIC REPORT.
FIFTH YEAR'S ANALYSIS. R. J. Newall and L. S.
Donian. 6p. (Forest Products Research Lao.
Supersedes corresponding part of Progress Report
711

This investigation consists in storing samples of ply-
wood and other glued wood products in selected loca-
tions for periods up to 10 years. At Intervals,
samples are removed and systematically tested for
deterioration of the glue lines, iungal attack, etc.
Inspections have been made at six-monthly intervals
over the past 5 years and a summary of the observa-
tions is presented.


N-38782'

Forest Products Research Lab. (Gt. Brit.)
COMPOSITE WOOD SECTION. TRIALS OF
TIMBERS FOR PLYWOOD MANUFACTURE.
ANINGUERIA ANINGUERIS ALTISSIMA UGANDA.
(NO RELIABLE WEIGHT FIGURES AVAILABLE BUT
PROBABLY BETWEEN 35 AND 40 LB. PER CUBIC
FOOT AT 15 PER CENT MOISTURE CONTENT).
(PROGRESS REPORT TWENTY-EIGHT). June 1955.
12p. taos. (Forest Products Research Lab.)



N-38783'

Forest Products Research Lao. (Gt. Brit.)
COMPOSITE WOOD SECTION. TRIALS OF
TIMBERS FOR PLYWOOD MANUFACTURE. ABURA
(NZINGU)-MITRAGYNA STIPULOSA UGANDA.
(36 POUNDS PER CUBIC FOOT AT 15 PER CENT
MOISTURE CONTENT). (PROGRESS REPORT
TWENTY-SEVEN). June 1955. 14p. tabs. (Forest
Products Research Lab.)



N-38784'

Forest Products Research Lab. (Gt. Brit.)
COMPOSITE WOOD SECTION. TRIALS OF
TIMBERS FOR PLYWOOD MANUFACTURE.
DAHOMA PIPTADENIA AFRICANA UGANDA.
147 POUNDS PER CUBIC FOOT AT 15 PER CENT
CONTENT). MUCHENCHE PIPTADENIA
BUCHANANII UGANDA. (35 POUNDS PER CUBIC
FOOT AT 15 PER CENT MOISTURE CONTENT).
(PROGRESS REPORT TWENTY-SIX). June 1955.
lip. tabs. (Forest Products Research Lab.)


N-38785*

Forest Products Research Lab. (Gt. Brit.)
MOISTURE RELATIONS OF COMPOSITE WOOD
PRODUCTS. PROGRESS REPORT TWENTY-
SEVEN JUNE 1955. THE FURROWING OF
VENEERED BLOCKBOARD. J. F. S. Carruthers.
9p. diagrs., tabs. (Forest Products Research Lab.
Supersedes Progress Report 26, May, 1954)





UNIVERSITY OF FLORIDA


3 1262 08153 278 9


The purpose of this investigation was to determine
the cause of the furrowing which sometimes occurs
on the surface of veneered blockboard after polishing.
Three different core constructions were employed
and an explanation of the furrowing is given for each
type.



N-38807*

Royal Aircraft Establishment (Gt. Brit.)
VELOCITY CALCULATIONS BY CONFORMAL MAP-
PING FOR TWO-DIMENSIONAL AEROFOILS.
D. A. Spence and N. A. Routledge. February 1955.
48p. diagrs., tabs. (RAE Aero 2539)

A method is derived for computing the conformal
transformation between the plane of an airfoil of
arbitrary shape (symmetrical or cambered), and the
plane of its velocity potential at zero lift (in which
the airfoil contour becomes a slit), in order to per-
mit calculations of the velocity at points off the sur-
face. The integral equation which relates the con-
tours is derived by an application of Cauchy's
theorem, and solved to the order of the square of
thickness ratio. The solution is found by repre-
senting the ordinate distribution by a Fourier series.
The rapid tailing-off of the Fourier coefficients for
all smooth airfoil shapes then leads to high accuracy
being achieved with a comparatively small amount of
effort. The method is straightforward and has
proved easy to use.



N-38808*

Royal Aircraft Establishment (Gt. Brit.)
THE CHARACTERISTIC FREQUENCIES OF SMALL
OSCILLATIONS IN THE FLOW PAST BLUFF
BODIES. D. A. Spence. May 1955. 23p. diagrs.
(RAE Aero 2532)

Summary: When a bluff body is placed in a steady
stream it experiences buffeting, the periodicity of
which can be explained in terms of interactions
between external and boundary layer regions. It
is shown that the frequency must satisfy a character-
istic equation in order for the oscillations induced in
the boundary layer to be compatible with those in the
outside stream. The equation is derived formally
for Lighthill's step case and for that of the circular
cylinder. The Karman vortices which are observed
in the latter case appear to be a consequence of the
oscillatory character of the circulation around the
cylinder.




DECLASSIFIED NACA REPORTS




NACA RM A54F28

ON THE RANGE OF APPLICABILITY OF THE
TRANSONIC AREA RULE. John R. Spreiter.
August 1954. 21p. (NACA RM A54F28) (Declassi-
fied from Confidential, 9/7/55)


NACA
RESEARCH ABSTRACTS NO. 90


Some insight into the range of appbcability of the
transonic area rule has been gained by comparison
with the appropriate similarity rule of transonic
flow theory and with experimental data for a large
family of rectangular wings having NACA 63AXXX
profiles.




NACA RM A54J07

THEORETICAL PRESSURE DISTRIBUTIONS FOR
SOME SLENDER WING-BODY COMBINATIONS AT
ZERO LIFT. Paul F. Byrd. January 1955. 39p.
diagrs. (NACA RM A54J07) (Declassified from
Confidential, 9/7/55)

Theoretical calculations are made of the pressure
distributions for some slender, symmetrical wing-
body combinations in subsonic and supersonic flow.
The combinations consist first of nonlifting, swept-
back wings mounted on a circular cylinder and
second of such wings mounted on a body indented so
that the local cross-sectional area of the combina-
tion is constant. The results indicate that indenta-
tion straightens out the isobars along the wing and
diminishes the maximum perturbation velocities.





NACA RM L52HO8

A STUDY OF THE ZERO-LIFT DRAG-RISE CHAR-
ACTERISTICS OF WING-BODY COMBINATIONS
NEAR THE SPEED OF SOUND. Richard T.
Whitcomb. September 1952. 41p. diagrs., photos.,
3 tabs. (NACA RM L52H08) (Declassified from
Confidential, 7/26/55)

Results are presented which indicate that near the
speed of sound the zero-lift drag rise of a thin low-
aspect-ratio wing-body combination is primarily de-
pendent on the axial distribution of the cross-
sectional areas normal to the airstream. Results
of an investigation of applications of this concept to
the reduction of the drag-rise increments of repre-
sentative wing-body combinations are also presented.





NACA RM L54A29a

ON SLENDER-BODY THEORY AT TRANSONIC
SPEEDS. Keith C. Harder and E. B. Kunker.
March 1954. 12p. (NACA RM L54A29a)
(Declassified from Confidential, 9/7/55)

The basic ideas of the slender-body approximation
have been applied to the nonlinear transonic-flow
equation for the velocity potential in order to obtain
some of the essential features of slender-body theory
at I ransonic speeds. The results of the investigation
are presented from a unified point of view which
demonstrates the similarity of slender-body solu-
tions in the various Mach number ranges. The tran-
sonic area rule and some conditions concerning its
validity follow from the analysis.


NACA Langley Field, Va.




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