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

Subjects

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:
AA00009235:00020

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N


National Advisory Committee For Aeronautics



Research Abstracts
3.59 I ARCH1 9- 1954


CURRENT NACA REPORTS

NACA Rept. 1123

A STUDY OF INVISCID FLOW ABOUT AIRFOILS AT
HIGH SUPERSONIC SPEEDS. A. J. Eggers, Jr.,
Clarence A. Syvertson and Samuel Kraus. 1953. n.
27p. diagrs., 6 labs. (NACA Rept. 1123. Formerly
TN 2646; TN 2729)

Steady flow about curved airfoils is investigated
analytically, first assuming air behaves as an ideal
gas, and then assuming it behaves as a thermally
perfect, calorically imperfect gas. Conclusions are
drawn from the study.


NACA TN 3063

EFFECTS OF WING POSITION AND FUSELAGE
SIZE ON THE LOW-SPEED STATIC AND ROLLING
STABILITY CHARACTERISTICS OF A DELTA
WING MODEL. Alex Goodman and David F.
Thomas, Jr. February 1954. 66p. diagrs., photos.,
3 tabs. (NACA TN 3063)

Results are presented of an investigation made to
determine the effects of wing position and fuselage
size on the low-speed static and rolling stability
characteristics of models having a triangular wing
and vertical tail surfaces. Interference increments
between the various components are evaluated, and
the variation of the vertical-tail lilt-curve slopes
and the efficiency factors with angle of attack as af-
fected by wing position and body size are presented.
TufI-grid pictures of the flow at the vertical tail as
affected by wing-fuselage interference are also pre-
sented.


NACA TN 3106

AN EVALUATION OF THE SOAP-BUBBLE METHOD
FOR BURNING VELOCITY MEASUREMENTS USING
ETHYLENE-OXYGEN-NITROGEN AND METHANE-
OXYGEN-NITROGEN MIXTURES. Dorothy M.
Simon and Edgar L. Wong. February 1954. 30p.
diagrs., photos., 5 tabs. (NACA TN 3106)

A nonaqueous soap-bubble method was used to meas-
ure the burning velocities of some ethylene-oxygen-
nitrogen and methane-oxygen-nitrogen mixtures.
Burning velocity calculations were based on high-
speed schlieren motion-picture records of the flame
growth and a theoretical expansion ratio. An upper
limit in the spatial velocity in the range 2500 to
3500 centimeters per second due to the appearance of
rough flames was found for the soap-bubble method.
Soap-bubble burning velocity measurements were
compared with measurements by other methods.


NACA RM 54A04 ..

EFFECTS OF MOLECJLARWWEIGHT ON CRAZING
AND TENSILE PROPERTIES OF POLYMETHYL
METHACRYLATE. I. Wolock, M. A. Sherman
and B. M. Axilrod, National Bureau of Standards.
February 1954. lip. diagrs., tab. (NACA
RM 54A04)

Tensile and crazing properties are reported for five
cast polymethyl-methacrylate sheets in which the
molecular weight of the resin was 90,000, 120,000.
200,000, 490,000. and 3,160,000. respectively.
Both stress and stress-solvent crazing tests were
conducted. The tensile strength and ultimate elon-
gation were found to increase rapidly with increasing
molecular weight at the lower molecular weights and
to level off at molecular weights of approximately
200,000 and 500,000. respectively. There was no
change in the modulus of elasticity over the range
studied. The specimens with lowest molecular
weight broke at low strains without crazing. The,--,,
stress and strain at which crazing occurred
creased with increasing molecular weighWs.j


9


BRITISH REPORTS



N-28467


/-


Aeronautical Research Council (Gt. Brit.)
KINETIC TEMPERATURE OF WET SURFACES. A
METHOD OF CALCULATING THE AMOUNT OF
ALCOHOL REQUIRED TO PREVENT ICE, AND THE
DERIVATION OF THE PSYCHROMETRIC EQUA-
TION. J. K. Hardy. 1953. 13p. 2 tabs. (ARC
R .i M 2830. Formerly NACA WR A-8; NACA
ARR 5G131

A method is given for calculating the temperature of
a surface wetted either by a pure liquid, such as
water, or by a mixture, such as alcohol and vater.
The method is applied to the problem of protecting.
by alcohol, propellers and the induction system of
the engine against ice. The minimum quantity of
alcohol required is calculated for a number of
arbitrarily chosen conditions. The effect of evapora-
tion of alcohol is shown by repeating the calculations
for a nonvolatile fluid. The method can be applied to
other problems in evaporation, for instance, to the
evaporation of fuel in the induction system ol the en-
gine. The psychrometric equation, used in wet-
bulb hygrometry, is deduced in its general form.
The effect of kinetic heating is included in this
equation.


* 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



U S1-ri-


/
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N-28468*

Aeronautical Research Council (Gt. Brit.)
OBSERVATIONS ON A THIN CAMBERED AERO-
FOIL BEYOND THE CRITICAL MACH NUMBER.
E. W. E. Rogers. (Read before VIth International
Congress of Applied Mechanics, September 1948)
1953. 16p. diagrs., photos. (ARC R & M 2432.
Formerly ARC 13,238; Perf. 685; FM 1455)

In the course of surface pressure measurements and
wake traverses on an airfoil section of 10 percent
thickness chord ratio, tested at high subsonic speeds
in the 20- by 8-inch (50.8 x 20.3 cm) high-speed wind
tunnel of the National Physical Laboratory, it was
discovered that at a particular incidence (3.70) an
extensive region of supersonic velocity (M = 1.15)
existed without the formation of a well-defined shock
wave or a rise of drag. The drag coefficient, in fact,
decreased markedly as the Mach number was raised
from a low value, and this was accompanied by a
rearward movement on the upper surface of the posi-
tion of boundary-layer transition corresponding to a
favorable change of the surface pressure gradient.
The transition positions measured with the "china-
clay" method are compared with those estimated
from the observed drag coefficients. Direct-
shadow photographs illustrate the development of the
shock-wave pattern.



N-28469*

Aeronautical Research Council (Gt. Brit.)
A GENERAL TREATMENT OF STATIC LONGITU-
DINAL STABILITY WITH PROPELLERS, WITH
APPLICATION TO SINGLE-ENGINED AIRCRAFT.
E. Priestley. 1953. 20p. diagrs., 2 tabs. (ARC
R & M 2732; ARC 7974. Formerly RAE Aero 1944)

A general method of treatment of stick-fixed static
longitudinal stability with propellers is given, distor-
tion and compressibility effects being neglected.
Model full-throttle data on some single-engined
fighters are analyzed for the flaps-up condition to
establish a basis of estimation of effect of propeller
on stability for this type of design. The general
effect of propellers on maneuver point, more
particularly the effect on Hm Kn, is considered in
an appendix.




N-28470*

Aeronautical Research Council (Gt. Brit.)
LOW-SPEED MEASUREMENTS OF THE PRESSURE
DISTRIBUTION AT THE SURFACE OF A SWEPT-
BACK WING. V. M. Falkner and Doris E. Lehrian.
1953. 36p. diagrs., photo., 27 tabs. (ARC
R & M 2741. Formerly ARC 12,647; Perf. 588;
S & C 2335)

Measurements were made at selected stations on a
sweplback wing with and without body. The wing has
450 sweepback, aspect ratio of 3, and constant chord.
The section was chosen to be suitable for work at low
Reynolds number. This is the first part of a pro-
gram on a sweptback wing to provide methods of cal-
culating the pressure at the surface, as well as the
general properties of wings. This report covers


NACA
RESEARCH ABSTRACTS NO. 59

what is thought to be one of the worst types of dis-
continuity likely to require investigation (excluding
deflected flaps); that is, sudden changes in direction
of both leading and trailing edges of the wing.

**

N-28471*

Aeronautical Research Council (GI. Brit.)
INTERIM REPORT ON V-G RECORDS ON HELI-
COPTERS. H. 1. Birds. 1953. 5p. diagrs.,
photos. (ARC R ., M 2746; ARC 12,327. Formerly
AFEE Rept. Rota 5)

V-g records have been obtained during the past year
on Hoverfly I helicopters. Some data have also been
obtained on a Hoverfly 11 and a Sikorsky S.51. The
V-g records on these aircraft were obtained mainly
during test flying which included blind flying and
some general flying. It was not possible to separate
the flight accelerations from the landing accelera-
tions, but these were small except in the case of
engine-off landings which were the subject of
separate tests.




N-28472*

Aeronautical Research Council (Gt. Brit.)
THE INITIAL BUCKLING OF A LONG AND
SLIGHTLY BOWED PANEL UNDER COMBINED
SHEAR AND NORMAL PRESSURE. E. H. Brown
and H. G. Hopkins. 1953. 19p. diagrs., 4 tabs.
(ARC R & M 2766; ARC 12,605. Formerly RAE
Structures 42)

Recent American experimental work has suggested
that the resistance to buckling of wing skin panels
under compression or shear loads is improved by
aerodynamic suction. A complete theoretical analy-
sis of this problem is very difficult, because, com-
pression load necessarily involves the consideration
of post-buckling behavior. An approach is made in
this report by considering the restricted problem of
the initial buckling of a long, thin, and slightly bowed
panel under combined shear and normal pressure.
The theoretical values of the initial shear buckling
stress which agree well with American experimen-
tal values, increase with both pressure and curva-
ture; the wave length of the buckles also increases
with pressure, but decreases with curvature. The
difference between the buckling stresses for simply
supported and clamped edges is considerable for a
flat panel under shear alone but decreases rapidly
with curvature and pressure, thus making the
indeterminacy of practical edge conditions of less
importance.





N-28473'

Aeronautical Research Council (Gt. Brit.)
TESTS ON A SWEPT-BACK WING AND BODY IN
THE COMPRESSED AIR TUNNEL. C. Salter, C. J.
W. Miles and H. M. Lee. 1953. 15p. diagrs.,
photos., 5 tabs. (ARC R & M 2738. Formerly ARC
13,155; Perf. 671)







NACA
RESEARCH ABSTRACTS NO. 59


The model was a sweplback wing of symmetrical
section and a long cylindrical body. Aspect ratio of
2.39, taper ratio of 0.37, and sweepback of the
quarter-chord of 42.50 gives the plan form of the
wing. The wing was 8.6 percent thick at the root and
10 percent thick at the tip. Results are given of the
lift, drag. and pitching moment for angles of attack
up to 300. A selection of How pictures is
reproduced.




N-28474*

Aeronautical Research Council (Gt. Brit.)
TWO-DIMENSIONAL AEROFOIL DESIGN IN COM-
PRESSIBLE FLOW. L. C. Woods. 1953. 19p.
diagrs., 3 tabs. (ARC R M 2731. Formerly
ARC 12,922; FM 1412)

This paper deals with the following two-dimensional
problem. "The design ol an airfoil to give a
specified velocity against chord curve at a given
free-stream Mach number." A "relaxation" method
is adopted, based on the differential equations for
incompressible and compressible flow. An essential
feature of the method is that the calculations are
carried out in the (0,*) or w-plane in which the air-
foil is represented by a slit along = 0. The square
mesh in this plane is formed by the streamlines
(4 = constant), and equipotentials (0 =-constant) for
incompressible flow about the airfoil. The method
is developed for a symmetrical airfoil at zero inci-
dence, but the modifications necessary for the more
general case are indicated. A worked example is
given, from which some idea of the accuracy of the
method can be gained. The compressible velocity
distribution about a known airfoil was taken as the
inirutial data. This airfoil was actually 12 percent
thick at 30 percent of the chord distance from the
leading edge. Using a mesh giving only 14 mesh
points on the airfoil, we find that the calculations
yield a 12.06 percent airfoil at 28.2 percent of the
chord distance from the leading edge.




N-28475*

Aeronautical Research Council (Gt. Brit.)
TESTS ON A WHIRLWIND AIRCRAFT IN THE
ROYAL AIRCRAFT ESTABLISHMENT 24-FT WIND
TUNNEL. (Includes: MOMENTUM INVESTIGA-
TIONS ON FUSELAGE-WING INTERFERENCE AND
NACELLE DRAG) T. V. Somer'.ille, R. R. Duddy
and G. H. L. Buxton. 1953. 17p. diagrs., 8 tabs.
(ARC R & M 2603. Formerly RAE BA Dept. Note
LWT 30; RAE BA Dept. Note LWT 34)

Simple modifications were found to decrease the drag
of the airplane. The drag analysis is not complete
and is focused chiefly on the drag due to leaks,
cooling, and excrescences. A complete record of the
tests is given. Modifications which gave an appreci-
able saving in drag were sealing of leaks and gaps,
fairing of exhaust cooling ducts, and fairing of main
cooling inlet. The saving in drag corresponds to an
increase in maximum speed of about 15 mph. A
further saving of 0.8 pound can be obtained by sealing
the cartridge chutes.


N-28476'

Aeronautical Research Council (Gt. Brit.)
THE INDUCED VELOCITY FIELD OF A ROTOR.
K. W. Mangler and H. B. Squire. 1953. 16p.
diagrs. (ARC R & M 2642; ARC 11,562; ARC
11,694. Formerly RAE Aero 2247; RAE Tech. Note
Aero 1958)

A short account and the results of a theoretical in-
vestigation of the velocity field induced by a lifting
rotor are given. The computation is based on the
assumptions that the rotor is lightly loaded and that
is has an infinite number of blades. This is applied
to calculate the induced velocity distribution for disk
incidences of 00, 150, 300, 450, and 900. For the
downwash at the rotor itself .(the normal component
of the induced velocity) the Fourier coefficients are
given, as they are needed for the calculation of the
blade motion.


N-28477*

Aeronautical Research Council (Gt. Brit.)
MEASUREMENTS OF THE AERODYNAMIC
DERIVATIVES FOR A HORN-BALANCED
ELEVATOR. N. C. Lambourne, A. Chinneck and
D. B. Betts. 1953. 15p. diagrs., photos., 2 tabs.
(ARC R & M 2653. Formerly ARC 12,085; 0.796)

This report gives the results of measurements by a
forced oscillation method of the direct derivatives
(aerodynamic stiffness and damping) for a horn-
balanced elevator. The tests were made at low air-
speeds on a complete wing-fuselage-tail model at 00
and 100 incidence in a wind tunnel. Some informa-
tion was obtained on the effect of mean elevator
angle on the derivatives when the model was at the
high incidence. Measurements were also made with
trailing-edge cords and transition wires in position.
The experiments suggest that none of the above fac-
tors causes a reduction in damping, but the stiffness
derivative was found to be considerably influenced by
the elevator angle and by the presence of trailing-
edge cords and transition wires. In general, the
measured values are numerically considerably less
than those calculated by simple strip theory using
two-dimensional vortex sheet theory results.



N-28478*

Aeronautical Research Council (Gt. Brit.)
MULTIPLE-JET WHITE-SMOKE GENERATORS.
C. Salter. 1953. 18p. diagrs., photos. (ARC
R & M 2657. Formerly ARC 10,296; FM 1058;
ARC 13,004; FM 1425)

Descriptions are given of equipment devised for the
generation of fairly large quantities of an optically
dense white smoke and special attention has been
paid to the need for delivering this through long
ducts or against an appreciable back-pressure. The
smoke consists of very small particles of condensed
paraffin vapor and is obtained by directing jets of
cool air on to high-speed jets of the vapor issuing
from very small orifices. The optimum outputs are
about 6 cu ft (170 litres) and 8 cu ft (230 litres) per
minute from No. 1 and No. 2 generators, respectively,
but considerably larger quantities can be delivered







4


with a slight loss of opacity. Under normal operating
conditions, the rate of use of paraffin is rather less
than 2 cu in. (33 c.c.) (No. 1) and 3 cu in. (49 c.c.)
(No. 21 per minute.



N-28549*

Aeronautical Research Council (Gt. Brit.)
THE MEASUREMENT OF HEAT TRANSFER AND
SKIN FRICTION AT SUPERSONIC SPEEDS. PART
IV. TESTS ON A FLAT PLATE AT M = 2.82. R. J.
Monaghan and J. R. Cooke. 1953. 42p. diagrs., 3
tabs. (ARC CP 140)

This note gives the results of overall heat transfer
and boundary-layer measurements made on a flat
plate in a 5 inch square supersonic wind tunnel
operating at M = 2.82 under atmospheric stagnation
pressure conditions. The tests were made to extend
the range of results previously obtained at M = 2.43
and used the same experimental equipment. In
general, the results confirm those obtained at the
lower Mach number and some general conclusions
are now drawn concerning the structure and behavior
of experimental laminar and turbulent compressible
boundary layers on a flat plate.



N-28550*

Aeronautical Research Council (Gt. Brit.)
CURVES FOR ESTIMATING THE WAVE DRAG OF
SOME BODIES OF REVOLUTION, BASED ON
EXACT AND APPROXIMATE THEORIES. L. E.
Fraenkel. 1953. 15p. diagrs. (ARC CP 136)

Curves are presented for estimating the wave drag,
at zero incidence, of forebodies and afterbodies
having straight and parabolic profiles. The after-
bodies are assumed to lie behind an infinitely long
cylindrical body. The curves are based on a limited
number of exact and second-order solutions which
have been generalized by appealing to the supersonic-
hypersonic similar ity law and to slender body and
quasi-cylinder solutions.



N-28551*

Aeronautical Research Council (Gt. Brit.)
DESIGN OF A RIGHT-ANGLED BEND WITH CON-
STANT VELOCITIES AT THE WALLS. A. S. Thom.
1953. 18p. diagrs., 5 tabs. (ARC CP 135)

In designing a corner in a two-dimensional duct, it
is possible, by the insertion of an airfoil, to main-
tain the same constant velocity on the outer and
inner walls. It is, however, necessary to shape
these walls to suit the conditions. The present
paper gives a method whereby the airfoil and walls
can be designed. Two examples are given.



N-28552*

Aeronautical Research Council (Gt. Brit.)
AIR COOLING METHODS FOR GAS TURBINE COM-
BUSTION SYSTEMS. F. J. Bayley. 1953. 44p.
diagrs. (ARC CP 133)


NACA
RESEARCH ABSTRACTS NO. 59

An account is given of the whole of the work which
has been done at NGTE on the problem of air-cooling
gas combustion systems. Each of the different
methods of wall cooling is discussed separately and
the theory and mechanism of the cooling process is,
developed from first principles. It is shown that
"sweat, or effusion cooling, is by far the most
effective and efficient method, while the use of
"louvered* surfaces represents the nearest practical
approach to this ideal which is possible while suita-
ble porous materials remain unavailable.


N-28553*

Aeronautical Research Council (Gi. Brit.)
THE MEASUREMENT OF HEAT TRANSFER AND
SKIN FRICTION AT SUPERSONIC SPEEDS. PART
II. MEASUREMENTS OF OVERALL HEAT TRANS-
FER AND OF THE ASSOCIATED BOUNDARY
LAYERS ON A FLAT PLATE AT MI = 2.43. R. J.
Monaghan and J. R. Cooke. 1953. 63p. diagrs..
3 tabs. (ARC CP 139)

A mean value of 0.906 was obtained for temperature
recovery factor and the overall heat transfer meas-
urements from plate to stream agreed well with re-
sults from the low-speed formula of reference 2.
There was some forward movement of boundary-
layer transition, a variation in the exponent of the
turbulent velocity distribution, and an increase in
displacement thickness with heat transfer. However,
no decrease in skin friction below its zero heat
transfer value was found.


N-28554*

Aeronautical Research Council (Gt. Brit.)
THE THEORY OF AEROFOILS WITH HINGED
FLAPS IN TWO-DIMENSIONAL COMPRESSIBLE
FLOW. L. C. Woods. 1953. 32p. diagrs. (ARC
CP 138)

Recently published methods of deducing practical
values of the various control characteristics from a
knowledge of their theoretical values increases the
importance of the theory of two-dimensional controls
in an inviscid compressible fluid. The classical
work of Glauert neglects compressibility and airfoil
thickness, and while the more recent work of
Goldstein and Preston includes thickness effects, it
ignores compressibility. Furthermore, this latter
method achieves accuracy for thick airfoils at the
cost of a complicated method of calculation. This
paper presents a theory of two-dimensional controls
in compressible flow which is almost as simple to
apply as Glauert's theory. An example given by
Goldstein and Preston is treated by the author's
method to illustrate this point.


N-28555*

Aeronautical Research Council (Gt. Brit.)
THE VIBRATIONS OF A SWEPT WING. N. S. Heaps.
1953. 26p. diagrs. (ARC CP 141)

The vibrations of a swept wing with ribs parallel to
the direction of flight are considered theoretically.
The couplings of torsion and flexure due to the
skewness of the ribs and the buiiding-in of the root
section are investigated.







NACA
RESEARCH ABSTRACTS NO. 59

N-28556'

Aeronautical Research Council (Gt. Brit.)
THE PERFORMANCE OF SOME TYPICAL TURBO-
JET ENGINE EXHAUST SYSTEMS, WITH PARTICU-
LAR REFERENCE TO THE EFFECTS OF SWIRL.
P. F. Ashwood and P. J. Fletcher. 1953. 28p.
diagrs., 3 tabs. (ARC CP 130)

Tests made to determine the effect of swirl on the
performance of a quarter-scale model of a typical
turbojet engine exhaust system and propelling nozzle
are described. The losses in the system were
derived from direct measurement of thrust. The
nondimensional thrust, expressed in terms of the
nozzle area and the total pressure at inlet to the ex-
haust diffuser, was found to vary linearly with the
ratio of ambient to inlet total pressures. An
annular nozzle was found to give slightly over 2 per-
cent more thrust at the choking condition than the
standard system for the same air mass flow.



UNPUBLISHED PAPERS


N-29228'

PROPOSED METHOD FOR DEFINING THE TAPER
RATIO OF MONOPLANE WINGS. (Una Proposta per
la determinazione del rapporto di rastremaxione
delle ali monoplane). Giuseppe Gabrielli. February
1954. 14p. diagrs. (Trans. from Onore di Modesto
Panetti, 1950, p.67-71)

A definition is given of taper ratio applicable to all
plan forms included within a double infinity of
analytically defined wings. A method of plotting the
plan form of a wing for whichh area, span, and taper
ratio values are given is described. The author de-
fines an equivalent wing to which static, aerodynamic,
aeroelastic, and weight calculations may be referred
in first approximation for any wing.




DECLASSIFIED REPORTS


NACA RM L6L26

FREE-FALL MEASUREMENTS AT TRANSONIC
VELOCITIES OF THE DRAG OF A WING-BODY
CONFIGURATION CONSISTING OF A 450 SWEPT-
BACK WING MOUNTED FORWARD OF THE MAXI-
MUM DIAMETER ON A BODY OF FINENESS RATIO
12. Charles W. Mathews and Jim Rogers Thompson.
April 2, 1947. 18p. diagrs., photo. (NACA
RM L6L26) (Declassified from Confidential,
11 10-53)

The National Advisory Committee for Aeronautics is
measuring drag of a series of complete airplane-like
configurations and their various components at tran-
sonic velocities by the free-fall method. This report
covers a test of one configuration of this series. The
configuration was composed of a 450 sweptback wing
of aspect ratio 4.1 mounted forward of the maximum
diameter of a 10-inch-diameter body of fineness ratio
12 equipped with stabilizing tail fins. The wing has a


5


70-inch span and incorporated an NACA 65-009 air-
foil section of 12-inch chord perpendicular to the
leading edge. The body-tail fin combination was
externally identical with a combination tested pre-
viously by this method.




NAGA RM L7C25a

TESTS OF A HORIZONTAL-TAIL MODEL THROUGH
THE TRANSONIC SPEED RANGE BY THE NACA
WING-FLOW METHOD. Richard E. Adams and
Norman S. Silsby. April 11, 1947. 24p. diagrs.,
photos., tab. (NACA RM L7C25a) (Declassified
from Restricted, 11/10/53)

A semispan model of the horizontal tail of a fighter
airplane was tested at transonic speeds. Measure-
ments of lift, elevator hinge moment, angle of
attack, and elevator angle were made in the Mach
number range from 0.75 to 1.04 for elevator de-
flections ranging from 100 to -10 and for angles of
attack of -1.20, 0.40, and 3.40. The hinge moment
data are considered to be only qualitative.




NACA RM L7D22

SOME PRESSURE-DISTRIBUTION MEASUREMENTS
ON A SWEPT WING AT TRANSONIC SPEEDS BY
THE NACA WING-FLOW METHOD. J. Ford
Johnston and Edward C. B. Danforth. June 6, 1947.
21p. diagrs., photos. (NACA RM L7D22) (Declassi-
fied from Restricted, 11/10/53)

First results are given of chordwise pressure-
distribution measurements on a 450 sweptback wing
at transonic speeds. These tests are part of a
fundamental investigate ion of flow phenomena near
sonic velocity by the NACA wing-flow method. Dis-
tributions were obtained at two spanwise extensions
of the half-span model of 2-inch chord and NACA
65-210 airfoil section measured perpendicular to the
leading edge. The aspect ratios were 2.1 and 3.5.


NACA-Langley 3-9-54 4M




UNIVERSITY OF FLORIDA

31 262 j 08 153 0174




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