Research abstracts

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Title:
Research abstracts
Physical Description:
93 v. : ; 27 cm.
Language:
English
Creator:
United States -- National Advisory Committee for Aeronautics
Publisher:
National Advisory Committee for Aeronautics
Place of Publication:
Washington, D.C
Publication Date:
Frequency:
irregular
completely irregular

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Subjects / Keywords:
Aeronautics -- Abstracts -- Periodicals   ( lcsh )
Aeronautics -- Research -- Abstracts -- Periodicals   ( lcsh )
Genre:
serial   ( sobekcm )
federal government publication   ( marcgt )
abstract or summary   ( marcgt )

Notes

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

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 001469326
notis - AGY1019
oclc - 01471285
lccn - 86657025
issn - 0499-9274
Classification:
lcc - TL501 .U5895
System ID:
AA00009235:00006

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National Advisory Committee for Aeronautics


Research Abstracts


NO.46


JULY 31, 1953


CURRENT NACA REPORTS

NACA Rept. 1095

TRANSONIC FLOW PAST A WEDGE PROFILE WITH
DETACHED BOW WAVE. Waller G. Vincenti and
Cleo B. Wagoner. 1952. ii, 30p. diagrs., tab.
(NACA Rept. 1095. Formerly TN 2339; TN 2588)

A theoretical study has been made of the aerodynamic
characteristics at zero angle of attack of a thin,
doubly symmetrical double-wedge profile in the range
of supersonic flight speed in which the bow wave is
detached. The analysis utilizes the equations of the
transonic small-disturbance theory and involves no
assumptions beyond those implicit in this theory.
The mixed flow about the front half of the profile is
calculated by relaxation solution of a boundary-value
problem for the transon4c small-dist rubance equation
in the hodograph plane (that is, the Tricomi equation)
The methods follow established lines except for the
somewhat novel treatment of the boundary conditions
along the shock polar and sonic line The purely
supersonic flow about the rear of the profile is found
by means of the method of characteristics
specialized to the transonic small-disturbance
theory. Complete calculations were made for four
values of the transoniruc similarity parameter. These
were found sufficient to bridge the gap between the
previous results of Guderley and Yoshihara at a Mach
number of 1 and the results which are readily ob-
tained when the bow wave is attached and the flow is
completely supersonic.

NACA TN 2966

PROPELLER-PERFORMANCE CHARTS FOR
TRANSPORT AIRPLANES. Jean Gilman, Jr.
July 1953 70p. diagrs (NACA TN 2966)

The preliminary selection of a propeller on the basis
of cruising and take-off performance for application
to transport airplanes at flight Mach numbers up to
0.8 can be accomplished by the use of the charts and
methods presented. The charts are of sufficient
scope to permit a fairly rapid evaluation of the pro-
peller performance for engine power ratings of
1.000 to 10,000 horsepower. The method is pre-
sented primarily in the interest of propeller-noise
abatement


NACA TN 2971

NT OF WATER DROPLETS ON
DIAMOND AIRFOILS AT SUPERSON-
C SPEED nrS. Seralini July 1953 62p.
O diagrs (NA 2971
**


An analytical solution has been obtained for the equa-
tions of motion of water droplets impinging on a
wedge in a two-dimensional supersonic low iield
with a shock wave attached to the wedge The
closed-form solution yields analytical expressions
for the equation of the droplet trajectory, the local
rate of impingement and the impingement velocity
at any point on the wedge surface, and the total rate
of impingement. The analytical expressions are
utilized to determine the inpingenment on the for-
ward surfaces of diamond airfoils in supersonic
flow fields with attached shock waves. The results
presented include the following conditions. droplet
diameters from 2 to 100 microns, pressure altitudes
from sea level to 30, 000 feet, free-streani static
temperatures from 4200 to -1600 R, free-stream
Mach numbers front 1 I to 2 0, semii-apex-aengltes,
for the wedge front 1 140 to 7.970, tckine to_ '-,
chord ratios for the diamond airfoy iro1 02 to .-
0. 14. chord lengths from 1 to 20 etandangles '
of attack from zero to the inverse tangent of the
airfoil thickness-to-chord ratio#, J 5

NACA TN 2972

THEORETICAL PRESSURE DIST RRUTION&ANb '
WAVE DRAGS FOR CONICAL BOATTALS. John
R. Jack July 1953 14p diagrs (NACA T1WN 2))

Afterbody pressure distributions and wave drags
were calculated using a second-order theory for a
variety of conical boattails at zero angle of attach
Results are presented for Mach numbers from 1 5
to 4. 5, area ratios front, 0 200 to 0 800, and boat -
tail angles froni 30 to 110 The results indicate that
for a given boattail angle. the wa\re drag decreases
witn increasing Mach number and area ratio The
wave drag, for a constant area ratio, increases with
increasing boattall angle For a specific Mach
number, area ratio, and fineness ratio. a compari-
son of the wave-uragf coefficients for conical.
tangent-parabolic, and secant-parabolic boattaiis
showed the conical boattail to have the smallest sate
drag.


NACA TN 2973

EFFECT OF PRESTRAINING ON RECRYSTALLILA-
TION TEMPERATURE AND MECHANICAL PROP-
ERTIES OF COMMERCIAL, SINTERED, WROUGHT
MOLYBDENUM Kenneth C. Dike and Roger A
Long July 1953. 25p. diagrs phios. 3 tabs
INACA TN 29731

Given three presumably identical lots of commercial,
sintered, wrought molybdenum, the 1-hour rtcrvs-
tallizatior' temperature ot one lot ren.ained above


f $gi ~ONLY
I DRESS REQUESTS DOCUMENTS TO NACA. 1794 F ST, NW, WASHINGTON 95 D C.CITING CODE NUMBER ABOVE EACH TITLE;
\T REPORT TITLE AUTHOR


^^_^>'^slr,







2

29000 F by limiting the amount of effective pre-
straining to 35 percent or less. Different recrys-
tallization temperatures were obtained in various
atmospheres, the highest in argon and the lowest
in hydrogen. Metal thus fabricated and then stress-
relieved possessed an ultimate tensile strength at
room temperature within 10 percent of metal swaged
99 percent and also possessed equivalent ductility.
At 18000 F, equivalent strength and ductility was ob-
tained irrespective of the amount of swaging over the
range of 10 to 99 percent. The amount of swaging
greatly influenced the recrystallized grain size but
the difference in grain size is not the major con-
trolling factor which determines whether recrystal-
lized molybdenum is ductile or brittle at room
temperature.

NACA TN 2974

EXPERIMENTS ON MIXED-FREE- AND -FORCED-
CONVECTIVE HEAT TRANSFER CONNECTED WITH
TURBULENT FLOW THROUGH A SHORT TUBE.
E. R. G. Eckert, Anthony J. Diaguila and Arthur N.
Curren. July 1953. 59p. diagrs., photo. (NACA
TN 2974)

Convective heat-transfer experiments were conducted
for turbulent, mixed-free- and -forced-convection
flow through a vertical heated tube with a length-to-
diameter ratio of 5. Studies were made with forced
air flowing parallel or opposite to the free convec-
tion flow. The Grashof numbers ranged from 109 to
1013 and the Reynolds numbers, from 36 x 103 to
377x103. These experiments together with results
obtained by other investigators revealed that the
total flow regime characterized by the Reynolds and
Grashof numbers can be divided into a forced-flow
regime, a free-flow regime, and a mixed-free- and
-forced-convection regime. The limits for the dif-
ferent regimes were established.

NACA TN 2975

STRUCTURAL EFFICIENCIES OF VARIOUS ALUMI-
NUM, TITANIUM,AND STEEL ALLOYS AT ELE-
VATED TEMPERATURES. George J. Heimerl and
Philip J. Hughes. July 1953. 16p. diagrs., tab.
(NACA TN 2975)

Efficient temperature ranges are indicated for two
high-strength aluminum alloys, two titanium alloys,
and three steels for some short-time compression-
loading applications at elevated temperatures. Only
the effects of constant temperatures and short ex-
posure to temperature are considered, and creep is
assumed not to be a factor. The structural efficien-
cy analysis is based upon preliminary results of
short-time elevated-temperature compressive
stress-strain tests of the materials. The analysis
covers strength under uniaxial compression, elastic
stiffness, column buckling, and the buckling of long
plates in compression or in shear.

NACA TN 2977

TECHNIQUES FOR CALCULATING PARAMETERS
OF NONLINEAR DYNAMIC SYSTEMS FROM RE-
SPONSE DATA. Benjamin R. Briggs and Arthur L.
Jones. July 1953. ii, 67p. diagrs., 4 tabs. (NACA
TN 2977)


NACA
RESEARCH ABSTRACTS'"-NO,46


A study has been made of the problem of determining
parameters of dynamic systems, such as aircraft or
servomechanisms, whose measured responses are
nonlinear. The parameters studied were functions
of displacement or velocity. The problem of deter-
mining such parameters involves both the recognition
of the degree of nonlinearity of the data and the
development of techniques of evaluation of the
parameters.

NACA TN 2981

THE HIGH-SPEED PLANING CHARACTERISTICS
OF A RECTANGULAR FLAT PLATE OVER A WIDE
RANGE OF TRIM AND WETTED LENGTH. Irving
Weinstein and Walter J. Kapryan. July 1953. 29p.
diagrs., photo., 2 tabs. INACA TN 2981)

The principal high-speed planing characteristics for
a prismatic surface having an angle of dead rise of
00 (flat bottom) have been determined over a wide
range of planing variables. Wetted length, resist -
ance, center-of-pressure location, and draft were
determined at speed coefficients ranging up to 25. 0,
beam loadings up to 87.3, and trims up to 300
Mean wetted lengths up to 7.0 beams were obtained
wherever possible.

NACA RM E53E04

MEASUREMENTS OF HEAT-TRANSFER AND FRIC-
TION COEFFICIENTS FOR AIR FLOWING IN A
TUBE OF LENGTH-DIAMETER RATIO OF 15 AT
HIGH SURFACE TEMPERATURES. Walter F.
Weiland and Warren H. Lowdermilk. July 1953.
19p. diagrs. (NACA RM E53E04)

Measurements of average heat-transfer and friction
factors were obtained for air flowing through a
smooth, electrically heated tube with a bellmouth
entrance and a length-to-diameter ratio of 15 for a
range of average surface temperature from 8750 to
17350 R and corresponding surface-to-bulk temp-
erature ratio from 1.6 to 2.8, Reynolds number
from 2200 to 300,000, and heat fluxes up to
230, 000 Btu/hr/ft2 of heat-transfer area. The data
of this investigation correlated with data obtained
in previous investigations with longer tubes on the
basis that the heat-transfer coefficient varies as the
-0. 1 power of the length-to-diameter ratio. No
effect of the length-to-diameter ratio was observed
on the average friction factor.

NACA RM E53E07

INVESTIGATION OF AERODYNAMIC AND ICING
CHARACTERISTICS OF A FLUSH ALTERNATE-
INLET INDUCTION-SYSTEM AIR SCOOP. James P.
Lewis. July 1953. 42p. diagrs., photos. (NACA
RM E53E07)

An investigation of the aerodynamic and icing charac-
teristics of a full-scale induction-system air-scoop
assembly incorporating a flush-type alternate inlet
was conducted in the NACA Lewis icing research
tunnel. The investigation was made over a range of
mass-air-flow ratios, angles of attack, airspeeds,
air temperatures, liquid-water content, and droplet
sizes. The ram inlet gave good pressure recovery





NACA
RESEARCH ABSTRACTS NO.46

in both clear air and icing, but rapid blocking of the
carburetor screen occurred in icing. The alternate
Inlet had poor pressure recovery in both clear air
and icing, but no serious screen icing was obtained.
The investigation included the use of preheat air
alone and in combination with ram- and alternate-
Inlet air.

NACA RM L52L29b

AN EXPERIMENTAL STUDY OF THE RELATION
BETWEEN AIRPLANE AND WIND-VANE MEASURE-
MENTS OF ATMOSPHERIC TURBULENCE. H. B.
Tolefson, K. G. Pratt and J. K. Thompson. July 1953.
20p. diagrs., photos., 3 tabs. (NACA RM L52L29b)

Analysis of simultaneous measurements of atmos-
pheric turbulence by wind vanes mounted on a tower
and by an airplane flying in the immediate vicinity of
the tower indicates good correlation between the gust
data for both sets of measurements. This correla-
tion indicates that surface and airplane measure-
ments of turbulence might be used interchangeably
for engineering and meteorological purposes.


NACA RM L53E05a

SOME MEASUREMENTS OF LANDING CONTACT
CONDITIONS OF TRANSPORT AIRPLANES IN
ROUTINE OPERATIONS. Norman S. Silsby,
Emanuel Rind and Garland J. Morris. July 1953.
8p. diagrs., photos. (NACA RM L53E05a)

Measurements have been obtained by means of a
specially built motion-picture camera of 126 land-
ings of current transport airplanes during routine
operations at the Washington National Airport in
clear-air daytime conditions. From these mea-
surements, sinking speeds, roll attitude angles, and
rolling velocities have been evaluated and a brief
statistical analysis of the results has been made.

NACA TM 1357

AIR ADMIXTURE TO EXHAUST JETS.
(Luftzumischung zu Abgasstrahlen). E. Singer.
July 1953. 35p. diagrs (NACA TM 1357. Trans.
from Ingenieur-Archiv, v. 18, no. 5, 1950, p. 310-
3231

The problem of thrust increase by air admixture to
exhaust jets of rockets, turbojet, ram- and pulse-
jet engines is investigated theoretically. The op-
timum ratio of mixing chamber pressure to ambient
pressure and speed range for thrust increase due to
air admixture is determined for each type of jet
engine.

NACA TM 1359

AIR-WATER ANALOGY AND THE STUDY OF HY-
DRAULIC MODELS. (La Similitudine Aria-
Acqua e lo Studio dei Modelli Idraalicln Giulio
Supino. July 1953. 22p. diagrs. (NACA TM 1359.
Trans. from Energia Elettrica, v. 28, no. 11, Nov.
1951)

The author first sets forth some observations about
the theory of models. Then he establishes certain


3

general criteria for the construction of dynamically
similar models in water and in air, through ref-
erence to the perfect fluid equations and to the ones
pertaining to viscous flow. It is, in addition,
pointed out that there are more cases in which the
analogy is possible than is commonly supposed.


BRITISH REPORTS


N-23173*

Department of Scientific and Industrial Research
(Gt. Brit.) EXPERIMENTS ON DISC-TESTING
MACHINES. (Bericht iiber dem stand der versuche
an den walzenpriifstiinden). 0. Kraupner. ii, 26p.
diagrs. (DSIR Sponsored Research (Germany)
Rept.2. Trans. from Technische Hochschule
Braunschweig, Institut fiir Maschinenelemente,
Einzelbericht 117, October 20, 1949)

The report presents experimental information on
materials, alinement, types of loading, lubricant
viscosity, rotational speed, and size of roller
bearings.


N-23175*

Forest Products Research Lab. (Gt. Brit.)
TRIALS OF TIMBERS FOR PLYWOOD MANU-
FACTURE. GABOON (AUCOUMEA KLAINEANA
PIERRE) FRENCH EQUATORIAL AFRICA. (27
POUNDS PER CUBIC FOOT AT 15 PER CENT
MOISTURE CONTENT) PROGRESS REPORT NINE-
TEEN. (Superseding Progress Report six,
February 1950) February 1953. 20p. 6 tabs.
(Forest Products Research Lab.)

This report presents the results of tests on gaboon,
a wood to be used in the manufacture of plywood.

N-23829*

Royal Aircraft Establishment (Gt. Brit.)
SOME ELECTRONIC MULTIPLIERS BASED ON
DIODE FUNCTION SHAPERS. 0. H. Lange and
G. J. Herring. April 1953. 35p. diagrs., photos.
(RAE Tech. Note GW 245)

Various types of analogue multipliers using elec-
tronic diodes are described in this report. They
are either of the indirect type based on the relation
xy = sin a sin 9 = cos(a 9) cos(a + ) or on the
2
parabolic law 4xy = (x + y)2 (x y)2 or direct
multipliers of a more general application. The
design procedure is given, and as examples some
indirect square law multipliers have been bull,
experimental results being presented. Figures ob-
tained from two different circuits showed an error
of less than 1 percent of the maximum value.


N-23830*

Aeroplane and Armament Experimental Establish-
ment (Gt. Brit.) DIFFERENTIAL PERFORMANCE
REDUCTION METHODS FOR TURBO-PROPELLER
AIRCRAFT. K. J. Lush. May 6, 1953. 16p.
2 tabs. (AAEE/Res/275)







4

Reduction methods for level speed, climb, and fuel
consumption tests have been derived, which are, in
the absence of compressibility effects, suitable for
use with turbopropeller aircraft at the present
stage. The methods require individual treatment of
the particular engine installation, as they deal only
with the reduction equation which would result from
any given engine behavior. Numerical generali-
zation about engine behavior did not seem worth
while at the present stage, but simplifying generali-
zations are briefly discussed, which might become
possible when experience is gained, and the simpli-
fying assumptions are indicated.

N-23831*

Royal Aircraft Establishment (Gt. Brit.)
PILOT ESCAPE FROM SPINNING AIRCRAFT.
T. H. Kerr. December 1952. 24p. photos.
(RAE Tech. Note Aero 2199)

A series of pilot escape tests from models of ele-
mentary and advanced trainers, and fighter aircraft
in the spin are presented. Escapes were made from
varying points relative to the wing chord, on the in-
board and outboard sides of the spin. The analyzed
results show that if the pilot requires to bail out
from a spinning aircraft, it is best to leave on the
outboard side of the aircraft and in the crouching
attitude. In this condition, it is most probable that
he will clear the aircraft cleanly and be outside the
spiral flight path within a half turn of the spin. If
he bails out on the inboard side, his flight path will
probably be through or very near the propeller disk
and it will probably take at least two turns of the
spin for him to clear the helical flight path of the
aircraft.

N-23865*

Aeronautical Research Council (Gt. Brit.)
THE SKIN FRICTION ON INFINITE CYLINDERS
MOVING PARALLEL TO THEIR LENGTH. G. K.
Batchelor. August 6, 1952. 15p. diagrs. (ARC
15, 105; FM 1772)

This paper is concerned with the three-dimensional
unidirectional motion that is generated by the forced
motion of an infinite cylinder parallel to its length in
a viscous incompressible fluid. The cylinder and
the fluid are at rest initially, and the cylinder is
brought to a velocity W which remains steady there-
after; the subsequent motion of the fluid is to be
determined.


UNPUBLISHED PAPERS
N-24151*

Massachusetts Inst. of Tech.
A RADIATION THERMOMETER FOR METEORO-
LOGICAL USE. J. C. Johnson. March 15, 1953.
43p. diagrs., photos., 3 tabs. (Massachusetts Inst.
of Tech.)
The theory, design, and construction of an infrared
pyrometer calibrated as a thermometer is described.
The instrument is termed a radiation thermometer.


NACA
RESEARCH ABSTRACTS'-NO 46 *

The radiation thermometer utilizes the 14.7A Dband of
radiation from the carbon dioxide always present in
air as an indicator of free air temperature. The
principle, though successful, averages the radiation
over a path of the order of 2 miles. Because of
this, the temperature measurement obtained will
have limited significance. The same instrument will
measure the temperature in the interior of clouds,
averaging over a path whose length is of the order of
the visual range. It is in this capacity that the
radiation thermometer may be most useful.

DECLASSIFIED NACA REPORTS


NACA RM A7K03

LOW-SPEED INVESTIGATION OF A SMALL TRI-
ANGULAR WING OF ASPECT RATIO 2. 0. I THE
EFFECT OF COMBINATION WITH A BODY OF
REVOLUTION AND HEIGHT ABOVE A GROUND
PLANE. Leonard M. Rose. August 27, 1948.
41p. diagrs., photos., 7 tabs. (NACA RM A7K03)
(Declassified from Restricted, 6/11/53)
Tests were made of triangular wing of aspect ratio
2. 0 with a symmetrical double-wedge section with
a maximum thickness of 5 percent of the chord.
Results are presented for the wing alone and in two
locations on a body of fineness ratio 12. 5. Further
data are also presented for the wing at several
heights above a ground plane.

NACA RM A7K05

TESTS OF A TRIANGULAR WING OF ASPECT RA-
TIO 2 IN THE AMES 12-FOOT PRESSURE WIND
TUNNEL. I THE EFFECT OF REYNOLDS NUM-
BER AND MACH NUMBER ON THE AERODYNAMIC
CHARACTERISTICS OF THE WING WITH FLAP
UNDEFLECTED. George G. Edwards and Jack D.
Stephenson. January 22, 1948. 42p. diagrs.,
photos. (NACA RM A7K05) (Declassified from Re-
stricted, 6/11/53)

This report presents results of tests of a semispan
model of a wing having a triangular plan form, an
aspect ratio of 2, and a double-wedge uncambered
airfoil section with the maximum thickness at 20 per-
cent of the chord. The effects of the addition of a
fuselage and of minor modifications to the airfoil
section were investigated. Aerodynamic character-
istics of the model are presented for a range of Rey-
nolds numbers from 5, 000, 000 to 27, 500, 000 at
0. 18 Mach number and for a range of Mach number
from 0. 18 to 0. 95 at 5, 300, 000 Reynolds number.
Change in Reynolds number had little effect, and
increasing the Mach number to the highest obtain-
able in the tunnel did not introduce sudden or a-
brupt changes in the wing characteristics.


NACA RM A7L05

AN INVESTIGATION OF THE DOWNWASH AND
WAKE BEHIND LARGE-SCALE SWEPT AND UN-
SWEPT WINGS. William H. Tolhurst, Jr.
February 2, 1948. 25p. diagrs., photo. (NACA
RM A7L05) (Declassified from Restricted, 6/11/53)






NACA
RESEARCH ABSTRACTS NO.46


This report contains the results of an investigation
to determine, at large scale, the downwash and wake
behind wings of 00, 1300, and *450 of sweep. The
results include contour maps of downwash angle in
a vertical plane located at a distance behind the wing
approximating a normal tail position, and the span-
wise variation of wake limits and pressure losses in
the wake for several angles of attack.

NACA RM A8D02

AERODYNAMIC STUDY OF A WING-FUSELAGE
COMBINATION EMPLOYING A WING SWEPT BACK
630 INVESTIGATION OF A LARGE-SCALE MODEL
AT LOW SPEED. Gerald M. McCormack and
Walter C. Walling. January 21, 1949. 20p. diagrs.,
photos. (NACA RM A8D02) (Declassified from
Restricted, 6/29 53)

Aerodynamic characteristics are presented for the
wing alone and for the wing-fuselage combination.
Experimental results are compared with character-
istics predicted theoretically. The relationships be-
tween the aerodynamic characteristics and flow con-
ditions existing over the wing are discussed.

NACA RM A8D27

WIND-TUNNEL INVESTIGATION OF TRANSONIC
AILERON FLUTTER OF A SEMISPAN WING MODEL
WITH AN NACA 23013 SECTION. Angelo Perone
and Albert L. Erickson. July 12, 1948. 23p.
diagrs., photos., tab. (NACA RM A8D27) (De-
classified from Confidential, 6/11/53)

An invest iat ion of aileron flutter with one degree of
freedom up to 0. 822 Mach number on a wing with an
NACA 23013 section and the effect of wing-tip relief
in alleviating aileron flutter are discussed. The
beneficial effect of a low aileron-span-to-wing-span
ratio is considered. The agreement that existed be-
tween the predicted and the actual flutter frequency
was sufficient to indicate that sections of the type
tested will not alter the existing recommendations
for flutter analysis.

NACA RM A8E03

TESTS OF A TRIANGULAR WING OF ASPECT
RATIO 2 IN THE AMES 12-FOOT PRESSURE WIND
TUNNEL. II THE EFFECTIVENESS AND HINGE
MOMENTS OF A CONSTANT-CHORD PLAIN FLAP.
Jack D. Stephenson and Arthur R. Amuendo.
September 21, 1948. 82p. diagrs., photos., 9 tabs.
(NACA RM A8E03) (Declassified from Restricted,
6/11 53)

This report presents the results of tests of a semi-
span model of a triangular wing having an aspect
ratio of 2. The effects of deflecting a constant-
chord plain flap were investigated for a Mach num-
ber range from 0. 18 to 0. 95 at a Reynolds number
of 5, 300, 000 and at low speed for a Reynolds num-
ber range from 5,300, 000 to 15, 000, 000. The data
have been used in the calculation of some longi-
tudinal characteristics of a triangular-wing tailless
airplane. Increasing the Mach number from 0. 18 to
0. 95 caused systematic increases in longitudinal
stability, lift, and control effectiveness, and alge-


5

braic decreases in the hinge-moment coefficients.
Little effect due to variation in Reynolds number or
to minor modifications of the airfoil section was
noted.

NACA RM A8H03

SOME PRELIMINARY RESULTS IN THE DETERMI-
NATION OF AERODYNAMIC DERIVATIVES OF CON-
TROL SURFACES IN THE TRANSONIC SPEED
RANGE BY- MEANS OF A FLUSH-TYPE ELECTRI-
CAL PRESSURE CELL. Albert L. Erickson and
Robert C. Robinson. October 8, 1948. 40p.
diagrs., photos., 3 tabs. (NACA RM A8H03) (De-
classified from Confidential, 6/11/53)

A method of measuring fluctuating aerodynamic
forces by using electrical pressure cells is pre-
sented along with the analysis of one cycle of motion
of a control surface during transonic flutter with one
degree of freedom. The report shows the differences
between upper- and lower-surface hinge moments
and explains these differences to some extent.


NACA RM A9B17

CHORDWISE AND SPANWISE LOADINGS MEA-
SURED AT LOW SPEED ON LARGE TRIANGULAR
WINGS. Adrien E. Anderson. April 19, 1949.
78p. diagrs., photos., 2 tabs. (NACA RM A9B17)
(Declassified from Restricted, 6/11/53)

Pressure distributions have been obtained from
three triangular wing models: a wing-alone model
having an aspect ratio of 2.04 and a modified double-
wedge airfoil section; the same wing combined with
a body of fineness ratio 12. 5; 'and a mock-up of a
triangular wing airplane which had an aspect ratio of
2.31 and an NACA 65-006. 5 airfoil section. Pres-
sure data were obtained through an angle-of-attack
range at zero angle of sideslip. Chordwise pres-
sure distribution, section lift characteristics, sec -
tion centers of pressure, and span load distribution
are presented.


NACA RM A9D04

FLIGHT INVESTIGATION OF THE EFFECT OF
BOUNDARY-LAYER SUCTION ON PROFILE-DRAG
COEFFICIENT AT SUPERCRITICAL MACH
NUMBERS. Richard B. Skoog. September 20, 1949.
30p. diagrs., photos. (NACA RM A9D04)
(Declassified from Confidential, 6/29/53)

Results of flight tests to study at high Reynolds num-
ber the effect of boundary-layer removal aft of the
shock wave on airfoil drag at supercritical Mach
numbers showed that no measurable effect of the
suction existed for the available suction coefficient
even though separation was present. The drag in-
crease with Mach number was due primarily to pres-
sure changes associated with supersonic flow on
upper and lower surfaces which resulted in increased
pressure drag.





6


NACA RM A9G13

TESTS OF A MODEL HORIZONTAL TAIL OF AS-
PECT RATIO 4.5 IN THE AMES 12-FOOT PRES-
SURE WIND TUNNEL. I QUARTER-CHORD LINE
SWEPT BACK 350. Bruce E. Tinling and Jerald K.
Dickson. September 9, 1949. 118p. diagrs ,
photo., tab. (NACA RM A9G13) (Declassified from
Restricted. 6/11/53)

This report presents results of tests of a semispan
model of a horizontal tail of aspect ratio 4. 5 with
the quarter-chord line swept back 350; the NACA
64A010 airfoil section; and a plain, sealed elevator
with tab. The Reynolds number was varied from
2, 000, 000 to 11, 000, 000 at a Mach number of 0. 21
and Mach number was varied from 0. 21 to 0.94 at a
Reynolds number of 2, 000, 000. Lift, drag, pitching
moment, elevator hinge moment, streamwise distri-
bution of static pressure at the mid-semispan, tab
effectiveness, tab hinge moment, and pressure dif-
ference across the elevator-nose seal were mea-
sured.


NACA RM A9G18

EFFECTS OF SYSTEMATIC CHANGES OF
TRAILING-EDGE ANGLE AND LEADING-EDGE
RADIUS ON THE VARIATION WITH MACH NUMBER
OF THE AERODYNAMIC CHARACTERISTICS OF A
10-PERCENT-CHORD-THICK NACA AIRFOIL SEC-
TION. James L. Summers and Donald J. Graham.
September 26, 1949. 81p. diagrs., photos., 9 tabs.
(NACA RM A9G18) (Declassified from Restricted,
6/11/53)

The effects of systematic changes in trailing-edge
angle and leading-edge radius on the variation with
Mach number of the aerodynamic characteristics of
a 10-percent-chord-thick symmetrical NACA air-
foil section are evaluated from wind-tunnel tests at
Mach numbers from 0. 3 to 0. 9. Reductions in both
parameters are shown to be beneficial to the section
lift-curve slope at Mach numbers below that for the
lift break, and to the maximum section lift coefficient
at Mach numbers above 0. 7 and detrimental to the
drag-divergence Mach number at low section lift
coefficients.

NACA RM A9H04

AN INVESTIGATION AT LOW SPEED OF A LARGE-
SCALE TRIANGULAR WING OF ASPECT RATIO
TWO. Ell. CHARACTERISTICS OF WING WITH
BODY AND VERTICAL TAIL. Adrien E. Anderson.
October 14, 1949. 96p. diagrs., photos., 3 tabs.
(NACA RM A9H04) (Declassified from Restricted,
6/11/53)

An investigation has been made to determine the
aerodynamic characteristics in sideslip of a tri-
angular wing of aspect ratio 2. 04 in combination
with a body of fineness ratio 12. 5 and a vertical
tail surface. Force and moment data were obtained
at several angles of sideslip for various configura-
tions of constant-chord split flaps, semispan split-
flap-type ailerons and a constant-chord rudder.
The Re-ynulds number was 15. 4 x 106 and the Mach
number 0 13


NACA
RESEARCH ABSTRACTS N0.46


NACA RM A9H1la


TESTS OF A MODEL HORIZONTAL TAIL OF AS-
PECT RATIO 4.5 IN THE AMES 12-FOOT PRES-
SURE WIND TUNNEL. II ELEVATOR HINGE LINE
NORMAL TO THE PLANE OF SYMMETRY. Bruce
E. Tinling and Jerald K. Dickson October 17,
1949. 106p. diagrs., photo., tabs. (NACA
RM A9H11la) (Declassified from Restricted,
6/11/53)

This report presents results of tests of a semispan
model of a horizontal tail of aspect ratio 4. 5 with
the hinge line normal to the plane of symmetry, the
NACA 64A010 airfoil section, and a plain, sealed
elevator with tab. The Reynolds number was varied
from 2, 000, 000 to 11,000,000 at a Mach number of
0. 21 and the Mach number was varied from 0. 21 to
0. 88 at a Reynolds number of 2, 000, 000. Lift,
drag, pitching moment, elevator hinge moment,
chordwise distribution of static pressure at the mid-
semispan, tab effectiveness, tab hinge moment, and
pressure difference across the elevator-nose seal
were measured. Results are compared to those
of a previous investigation of a model which dif-
fered primarily in that its quarter-chord line had
350 of sweepback.

NACA RM A9KO2a

EFFECTS OF BOUNDARY-LAYER CONTROL ON
THE LONGITUDINAL CHARACTERISTICS OF A 45
SWEPT-FORWARD WING-FUSELAGE COMBINA-
TION. Gerald M. McCormack and Woodrow L. Cook.
February 2, 1950. 73p. diagrs., photo., 2 tabs.
(NACA RM A9K02a) (Declassified from Restricted,
6/29/53)

An investigation has been conducted to determine the
benefits obtainable by applying boundary-layer con-
trol to a 450 sweptforward wing-fuselage combination.
Force and pressure-distribution data are presented
for the model with and without boundary -layer con-
trol with leading-edge and trailing-edge flaps. The
results showed that with suction applied through a
slot at the wing-fuselage juncture separation could be
postponed from a lift coefficient between 30 and 50
percent of the maximum to a lift coefficient between
78 and 93 percent of the maximum.

NACA RM A9KO02b

A DESIGN STUDY OF LEADING-EDGE INLETS FOR
UNSWEPT WINGS Robert E. Dannenberg.
June 30, 1950. 56p. diagrs., photos., 3 tabs.
(NACA RM A9K02b) !Declassified from Restricted,
6/11/53)

A method is presented for calculating the profile co-
ordinates for an inlet to be placed in the leading edge
of an airfoil from formulas. The method includes
an application of the principles of thin-airfoil theory
which permits the change in velocity distribution
caused by a variation in inlet profile to be calcu-
lated. Wind-tunnel tests of leading-edge inlets in an
airfoil having the NACA 631-012 section were made
to evaluate the effects of the inlets on the aerody-
namic characteristics of the airfoil. The results
indicated that the airfoil with an inlet devised by the
design method had satisfactory aerodynamic charac-
teristics


*i -






NACA
RESEARCH ABSTRACTS NO.46

NACA RM A50D04

EFFECTS OF LEADING-EDGE RADIUS AND MAX-
IMUM THICKNESS-CHORD RATIO ON THE VARIA-
TION WITH MACH NUMBER OF THE AERODYNAM-
IC CHARACTERISTICS OF SEVERAL THIN NACA
AIRFOIL SECTIONS Robert E. Berggren and
Donald J. Graham July 3, 1950. 65p. diagrs.,
7 tabs. INACA RM A50D04) (Declassified from
Restricted, 6 11 53)

The results of a wind-tunnel investigation at Mach
numbers to approximately 0.9 and Reynolds numbers
from 1 x 106 to 2 x 106 indicate no significant effects
of leading -edge-radius variation on the variation
with Mach number of the aerodynamic characteris-
tics of 4- and 6-percent-chord-thick NACA 4-digit-
series airfoil sections. The results indicate bene-
ficial effects of maximum thickness reduction on lift
and drag characteristics and no important effects
on moment characteristics.

NACA RM A50E02

PRESSURE-DISTRIBUTION AND RAM-RECOVERY
CHARACTERISTICS OF NACA SUBMERGED IN-
LETS AT HIGH SUBSONIC SPEEDS. Joseph L.
Frank. July 7, 1950. 41p. diagrs., photo.
(NACA RM A50E02) (Declassified from Confiden-
tial, 6/11/53)

Presented are results of tests of NACA submerged
inlets at four fuselage stations on a model of a hypo-
thetical fighter airplane. Ram-recovery ratios
were generally maximum for mass-flow ratios be-
tween 0.60 and 0. 80. The inlet in the most for-
ward position provided the highest ram recovery for
almost all test conditions. Inlets in the region of
high-velocity flow induced by the wing had compress-
ibility losses beginning at 0. 70 Mach number.

NACA RM E7K19

PRELIMINARY INVESTIGATION OF CONE-TYPE
DIFFUSERS DESIGNED FOR MINIMUM SPILLAGE
AT INLET. Roger W. Luidens and Henry Hunczak.
May 3, 1948. 31p. photos., diagrs. (NACA
RM E7K 19) (Declassified from Confidential,
6/11/53)

Presents an investigation at a Mach number of 1. 85
of cone-type diffuser designed for minimum spillage
at the inlet. Pressure recoveries of stationary-
cone configurations are presented as functions of
cone angle, outlet-inlet area ratio, throat length,
and angle of attack. Pressure recoveries of a 300
movable-cone configuration are presented as func-
tions of contraction ratio and outlet-inlet area ratio.
Stabilization of the normal shock within the diffuser
by use of perforations is experimentally verified.

NACA RM E7L05

ANALYTICAL INVESTIGATION OF DISTRIBUTION
OF CENTRIFUGAL STRESSES AND THEIR RE-
LATION TO LIMITING OPERATING TEMPERA-
TURES IN GAS-TURBINE BLADES. Richard H.
Kemp and William C. Morgan. April 12, 1948.
25p. photo., diagrs. (NACA RM E7L05) (De-
classified from Restricted, 6/11/53)


7


The distribution of centrifugal stresses at maximum
operating conditions are presented for six turbine-
blade designs and the limitation imposed on the per-
missible centrifugal stresses by the operating tem-
peratures is discussed. It was found that this tem-
perature limitation was important. In the case of
one blade, several recorded surface failures have
occurred in the region indicated as critical by this
analysis.

NACA RM E8A26

FREE-FLIGHT INVESTIGATION OF 16-INCH-
DIAMETER SUPERSONIC RAM-JET UNIT. George
F. Kinghorn and John H. Disher. May 28, 1948.
20p. diagrs., photos. (NACA RM E8A26) (De-
classified from Confidential, 6/11/53)

In order to determine the performance of ram-jet
units in flight at supersonic speeds and to study
methods for improving the performance, an investi-
gation of a series of 16-inch-diameter ram-jet en-
gines is being conducted in free flight. Supersonic
flight speeds are obtained by releasing the unit from
an airplane at an altitude of approximately 30, 000
feet and allowing the engine thrust and the force of
gravity to accelerate the unit.

NACA RM E8H03

COOLING OF GAS TURBINES. IX COOLING EF-
FECTS FROM USE OF CERAMIC COATINGS ON
WATER-COOLED TURBINE BLADES. W. Byron
Brown and John N. B. Livingood. October 13, 1948.
26p. diagrs. (NACA RM E8H03) (Declassified
from Restricted, 6/11/53)

Calculations of the effect of a ceramic coating on
blade-metal and gas temperatures are prepared and
comparisons of these effects with the changes pro-
duced by moving the cooling passages nearer the
trailing edge are made. Results indicate that very
low-conductivity ceramics, such as mica and por-
celain, or material with similar conductivities,
must be used to secure practical results.

NACA RM E8K05

PRESSURE DISTRIBUTIONS ON THIN CONICAL
BODY OF ELLIPTIC CROSS SECTION AT MACH
NUMBER 1. 89. Stephen H. Maslen. January 20,
1949. 18p. diagrs., photo. (NACA RM E8K05)
(Declassified from Confidential, 6/11/53)

Experimental pressure distributions at Mach number
1.89 are presented for a conical body of elliptic
cross section. Data are presented for range of an-
gles of yaw from -160 to 160 and angles of attack
from -100 to 100. Over the entire range of calcu-
lations, linearized theory checked closely with ex-
periment.

NACA RM E50B02

CHARACTERISTICS OF PERFORATED DIFFUSERS
AT FREE-STREAM MACH NUMBER 1. 90. Henry
R. Hunczak and Emil J. Kremzier. May 8, 1950.
69p. diagrs., photos. (NACA RM E50B02) (De-
classified from Confidential, 6/11/53)







a

An investigation was conducted at Mach number 1. 90
to determine pressure recovery and mass-flow char-
acteristics of series of perforated convergent-
divergent supersonic diffusers. Pressure re-
coveries as high as 96 percent were obtained, but at
reduced mass flows through the diffuser. Theoret-
ical considerations of effect of perforation distribu-
tion on shock stability in converging section of dif-
fuser are presented and correlated with experimen-
tal data. A method of estimating relative impor-
tance-of pressure recovery and mass flow on inter-
nal thrust coefficient basis is given and a comparison
of various diffusers investigated is made.

NACA RM E50H22

EFFECT OF RAM-JET PRESSURE PULSATIONS
ON SUPERSONIC-DIFFUSER PERFORMANCE.
James F. Connors. November 20, 1950. 29p.
photos., diagrs. (NACA RM E50H22) (Declassi-
fied from Confidential, 6/11/53)

Experimental study of effects of combustion rough-
ness and controlled mechanical oscillations on dif-
fuser operation was conducted on an 8-inch ram jet
at Mach number 1. 87. Within experimental accura-
cy, the optimum mean combustion-chamber static
pressure occurred when the maximum instantaneous
static pressure equalled the optimum steady-flow
value and the corresponding decrement in optimum
mean static pressure was equal to one-half the total
amplitude of the pressure pulsations. A marked
attenuation of amplitude of cold-buzz pressure fluc-
tuations was effected over limited range of subcrit-
igal diffuser operation by means of properly tuned
rotating disk located in combustion chamber.

NACA RM E50K07

AN INVESTIGATION OF CONVERGENT-DIVERGENT
DIFFUSERS AT MACH NUMBER 1. 85. DeMarquis
D. Wyatt and Henry R. Hunczak. February 2, 1951.
60p. diagrs., photos., tab. (NACA RM E50K07)
(Declassified from Restricted, 6/11/53)

Total-pressure recoveries and internal static-
pressure distributions for a series of convergent-
divergent supersonic diffusers operating at a Mach
number of 1. 85 are presented. The effect of in-
serting a cylindrical section between the supersonic
inlet and the subsonic diffuser on shock stability is
experimentally determined.


NACA RM E51B10

THE USE OF PERFORATED INLETS FOR EFFI-
CIENT SUPERSONIC DIFFUSION. John C. Evvard
and John W. Blakey. (Revised) April 13, 1951.
36p. diagrs., photo. (NACA RM E51B10) (De-
classified from Restricted, 6/11/53)

The use of wall perforations on supersonic diffusers
to avoid the internal contraction-ratio limitation is
described. Experimental results at a Mach number
of 1. 85 on a preliminary model of a perforated dif-
fuser having a geometric internal contraction ratio
of 1. 49 (the isentropic value) are presented. A
theoretical discussion of the flow coefficients as well
as the size and the spacing of the perforations is also


NACA
RESEARCH ABSTRACTS-'-0.46

included. At angles of attack of 00, 30, and 50,
total-pressure recoveries of 0. 931, 0.920. and
0.906, respectively, were obtained.

NACA RM E51B13

THEORETICAL AND EXPERIMENTAL INVESTIGA-
TION OF ADDITIVE DRAG. Merwin Sibulkin.
May 21, 1951. 33p. diagrs. (NACA RM E51BL3)
(Declassified from Confidential, 6/11/ 53)

The significance of additive drag is discussed and
equations for determining its approximate value are
derived. Charts are presented giving values of
additive drag for open-nose inlets and for annular-
nose inlets with conical flow at the inlet. The ef-
fects of variable inlet total-pressure recovery and
static pressures on the center body are investigated.
and an analytical method of predicting the variation
of pressure on the center body with mass-flow ratio
is given. Experimental values of additive drag are
compared with values predicted by the methods pre-
sented.

NACA RM L7G09

INVESTIGATION OF HIGH-LIFT AND STALL-
CONTROL DEVICES ON AN NACA 64-SERIES 420
SWEPTBACK WING WITH AND WITHOUT FUSE-
LAGE. Robert R. Graham and D. William Conner.
October 14, 1947. 47p. diagrs., photos., 2 tabs.
(NACA RM L7G09) (Declassified from Restricted,
6/25/53)

Contains low speed, high Reynolds number charac-
teristics of a 420 sweptback wing with aspect ratio 4
and NACA 64-series sections tested with various
combinations of high-lift and stall-control devices
and a fuselage. Combinations of outboard leading-
edge flaps or slats with inboard split flaps sub-
stantially increased the maximum lift coefficient
while maintaining longitudinally stable character-
istics at the stall. Neither upper-surface fences nor
outboard upper-surface split flaps were effective in
improving longitudinal stability for those combina-
tions where sizeable positive pitching-moment in-
creases occurred beyond the stall.

NACA RM L7J20

INITIAL FLIGHT TESTS OF THE NACA FR-2, A
HIGH-VELOCITY ROCKET-PROPELLED VEHICLE
FOR TRANSONIC FLUTTER RESEARCH. J. G.
Barmby and J. M. Teitelbaum. March 4, 1948.
21p. diagrs., photos. (NACA RM L7J20) (De-
classified from Restricted. 6/11/53)

The initial flight tests of two simplified flutter ve-
hicles, which were launched at the Langley Pilotless
Aircraft Research Station at Wallops Island, Va.,
are described in this report. The results of the
tests are in agreement with the results of the freely-
falling-body test in that the wing failures in the tran-
sonic range occurred at velocities greater than the
flutter velocity calculated from the two-dimensional,
incompressible theory. Although the simple bread-
wire system seems satisfactory for exploratory
tests, the scope of the investigation could be ex-






NACA
RESEARCH ABSTRACTS NO.46

tended by use of a frequency-recording telemeter.
The use of high-velocity-rocker test vehicles of this
type offers promise for flutter testing in the tran-
sonic and the supersonic regions.

NACA RM L7K06

TWO-DIMENSIONAL WIND-TUNNEL INVESTIGA-
TION AT HIGH REYNOLDS NUMBERS OF AN NACA
65A006 AIRFOIL WITH HIGH-LIFT DEVICES.
Robert J. Nuber and Stanley M. Gottlieb.
February 4, 1948. 28p. diagrs., photos., tab.
(NACA RM L7K06) (Declassified from Restricted,
6 '29 53)

An investigation was made of an NACA 65A006 air-
foil equipped with high-lift devices consisting of a
0.15-chord drooped-nose flap and a 0.20-chord plain
trailing-edge flap. The airfoil section lift, pitching-
moment, and drag characteristics, obtained at high
Reynolds numbers and low Mach numbers (M 0.14)
with the flaps deflected individually and simultane-
ously, are presented.

NACA RM L8D29

WIND-TUNNEL INVESTIGATION OF HIGH-LIFT
AND STALL-CONTROL DEVICES ON A 370 SWEPT-
BACK WING OF ASPECT RATIO 6 AT HIGH
REYNOLDS NUMBERS. William Koven and
Robert R. Graham. September 2, 1948. 70p.
diagrs., photos., tab. (NACA RM L8D29)
(Declassified from Restricted, 6/29/53)

Stall-control devices including an extensible round-
nose leading-edge flap, a leading-edge slat, and a
drooped leading edge were investigated; half- and
full-span trailing-edge split and double slotted flaps
were also investigated on the plain wing and in com-
bination with the leading-edge devices. Force and
moment data and stall studies are presented for all
configurations; downwash and dynamic-pressure
measurements are given for one combination of slat
and double slotted flap. Typical plots of variation of
spanwise position of load centroid with lift coeffi-
cient are included.

NACA RM L8E18

THE EFFECT OF BOUNDARY-LAYER CONTROL
BY SUCTION AND SEVERAL HIGH-LIFT DEVICES
ON THE LONGITUDINAL AERODYNAMIC CHARAC-
TERISTICS OF A 47.50 SWEPTBACK WING-
FUSELAGE COMBINATION. Jerome Pasamanick
and Anthony J. Proterra. November 4, 1948. 44p.
diagrs., photo., 2 tabs. (NACA RM L8E18)
(Declassified from Restricted, 6/29/53)

An investigation has been made in the Langley full-
scale tunnel of a 47.50 sweptback wing-fuselage
combination equipped for boundary-layer control by
suction. The wing aspect ratio was 3.5, the taper
ratio was 0.5, and the airfoil sections normal to the
quarter-chord line were NACA 641-A112. The wing
configurations tested included the wing with various
combinations of extensible leading-edge and split
flaps. The effect of Reynolds number, suction-slot
location, and suction flow coefficient on the aero-
dynamic characteristics was determined for the
model at zero yaw over a range of angle of attack.


9

NACA RM L8E21

THE EFFECT OF BOUNDARY-LAYER CONTROL
BY SUCTION AND OF SEVERAL HIGH-LIFT DE-
VICES ON THE AERODYNAMIC CHARACTERISTICS
IN YAW OF A 47.50 SWEPTBACK WING-FUSELAGE
COMBINATION. Jerome Pasamanick. October 28,
1948. 36p. diagrs., photo. (NACA RM L8E21)
(Declassified from Restricted, 6/29/53)

An investigation has been made in the Langley full-
scale tunnel of a 47.50 sweptback wing-fuselage
combination equipped for boundary-layer control by
suction. The wing aspect ratio was 3.5, the taper
ratio was 0.5, and the airfoil sections normal to the
quarter-chord line were NACA 641-A112. Tests in-
cluded the plain wing and the wing with various com-
binations of extensible leading-edge and split flaps.
The investigation was made to determine the effect
of boundary-layer control by suction on the aerody-
namic characteristics in yaw and on the effectiveness
of a split-flap-type aileron for a range of angle of
attack and suction-flow coefficient at a Reynolds
number of 4.2 x 106 corresponding to a Mach number
of approximately 0.07.

NACA RM L8J29

PRESENT STATUS OF RESEARCH ON BOUNDARY-
LAYER CONTROL. Albert E. von Doenhoff and
Laurence K. Loftin, Jr. January 12, 1949. 40p.
diagrs. (NACA RM L8J29) (Declassified from
Confidential, 6/29/53)

A survey has been made of the present status of
boundary-layer control and its applications in aero-
nautics. The possible improvements in airplane
characteristics resulting from boundary-layer con-
trol are discussed and the general lines of future
research are indicated.

NACA RM L9B04

HYDRODYNAMIC CHARACTERISTICS OF AERO-
DYNAMICALLY REFINED PLANING-TAIL HULLS.
Robert McKann and Claude W. Coffee. March 28,
1949. 41p. diagrs., photos. (NACA RM L9B04)
(Declassified from Confidential, 6/11/53)

The hydrodynamic characteristics of two aerody-
namically refined planing-tail hulls were determined
from dynamic model tests in Langley tank no. 2.
Stable take-offs could be made for a wide range of
location of the center of gravity. The lower por-
poising limit was high but no upper limit was en-
countered. The hump load-resistance ratios were
low. Stable landings were made at the center-of-
gravity locations tested for a wide range of contact
trim.


NACA RM L9D15

HYDRODYNAMIC CHARACTERISTICS OF A SWEPT
PLANING-TAIL HULL. Robert E. McKann, Claude
W. Coffee and Donald D. Arabian. September 12,
1949. 34p. diagrs., photos., tab. (NACA
RM L9D15) (Declassified from Confidential,
6/11/53)






10

The hydrodynamic characteristics of an aero-
dynamically refined swept planing-tail hull were
determined from dynamic model tests in Langley
tank no. 2. Acceptable propeller spray character-
istics were obtained by the use of vertical spray
strips. Stable take-offs could be made with a large
range of elevator deflections and center-of-gravity
location. Stable landings were made for a wide
range of center-of-gravity location and contact trim.
The hump load-resistance ratio was low.

NACA RM L9FO8

AERODYNAMIC CHARACTERISTICS OF AN
AIRFOIL-FOREBODY SWEPT FLYING-BOAT HULL
WITH A WING AND TAIL SWEPT BACK 51.30 AT
THE LEADING EDGE. Rodger L. Naeseth and
Richard G. MacLeod. September 9, 1949. 27p.
diagrs., photo., tab. (NACA RM L9F08)
(Declassified from Confidential, 6/11/53)

Results of tests in the Langley 300 MPH 7- by 10-
foot tunnel of an airfoil-forebody swept flying-boat
hull mounted on a wing swept back 51.30 at the
leading edge are presented. Aerodynamic charac-
teristics of the hull (including wing interference)
and longitudinal and lateral stability characteristics
of the complete configuration (hull, wing, 51.30
sweptback tail) with stall-control devices and lift
flaps are presented.

NACA RM L9G15

INVESTIGATION AT LARGE SCALE OF THE PRES-
SURE DISTRIBUTION AND FLOW PHENOMENA
OVER A WING WITH THE LEADING EDGE SWEPT
BACK 47.50 HAVING CIRCULAR-ARC AIRFOIL
SECTIONS AND EQUIPPED WITH DROOPED-NOSE
AND PLAIN FLAPS. Roy H. Lange, Edward F.
Whittle, Jr. and Marvin P. Fink. September 8, 1949.
72p. diagrs., 3 tabs. (NACA RM L9G15)
(Declassified from Restricted, 6/29/53)

The results of an investigation of the effects of
drooped-nose-flap and plain-flap deflection on the
pressure distribution over a wing with the leading
edge swept back 47.50 and having symmetrical
circular-arc airfoil sections conducted in the
Langley full-scale tunnel at a Reynolds number of
4.3 x 106 and a Mach number of 0.07 are included in
(his paper. The configurations tested include the
basic wing, the wing with a full-span drooped-nose
flap, an inboard semispan plain flap, and a combina-
tion of these two flap configurations.


NACA RM L9H10

INVESTIGATION OF HIGH-SUBSONIC PERFORM-
ANCE CHARACTERISTICS OF A 120 21-INCH
CONICAL DIFFUSER, INCLUDING THE EFFECTS
OF CHANGE IN INLET-BOUNDARY-LAYER
THICKNESS. Martin R. Copp and Paul L. Klevatt.
March 24, 1950. 51p. diagrs., photos. (NACA
RM L9H10) (Declassified from Restricted,
6/11/53)

An investigation of performance characteristics and
boundary-layer growth of a 120 21-inch conical
diffuser over a Reynolds number range of 1. 45 x


NACA
RESEARCH ABSTRACTS N0.46

106 to 7.45 x 106 and a Mach number range of 5 I. I.
to approximately choking are presented for two
thicknesses of inlet boundary layer. The mean
value, over the entire range of inlet velocities, of
the displacement thickness of the thinner inlet
boundary layer was approximately 0.035 inch and
that of the thicker inlet boundary layer was approx-
imately six times this value. Performance param-
eters, static-pressure distributions, and boundary-
layer measurements are presented for both thick-
nesses of inlet boundary laver.

NACA RM L9K10

THE EFFECT OF THE INLET MACH NUMBER AND
INLET-BOUNDARY-LAYER THICKNESS ON THE
PERFORMANCE OF A 230 CONICAL-DIFFUSER -
TAIL-PIPE COMBINATION. Jerome Persh.
March 21, 1950. 53p. diagrs. (NACA RM L9K10)
(Declassified from Restricted, 6/11/53)

An investigation was made of the effect of the inlet
Mach number and entrance-boundary-layer thickness
on the performance of a 230 21-inch conical-
diffuser tail-pipe combination with a 2:1 area ratio
over an inlet Mach number range of 0. 17 to 0. 89 and
corresponding Reynolds numbers of 1, 700, 000 to
7,070,000. Performance coefficients, longitudinal
static-pressure distributions, and the results of
boundary-layer surveys made at six stations along
the diffuser wall are presented.

NACA RM L50B15

FULL-SCALE INVESTIGATION OF BOUNDARY-
LAYER CONTROL BY SUCTION THROUGH
LEADING-EDGE SLOTS ON A WING-FUSELAGE
CONFIGURATION HAVING 47.50 LEADING-EDGE
SWEEP WITH AND WITHOUT FLAPS. Jerome
Pasamanick and Thomas B. Sellers. April 5, 1950.
55p. diagrs., photo., 2 tabs. (NACA RM L50B15)
(Declassified from Restricted, 6/29/53)

Presents results of the investigation in the Langley
full-scale tunnel of leading-edge boundary-layer con-
trol by suction in combination with high-lift devices
on a wing-fuselage configuration having 47.50 leading
edge sweep. The wing aspect ratio was 3.4, the
taper ratio was 0.51, and the airfoil sections normal
to the quarter-chord line were NACA 641A112. The
data include the effects of suction-slot location and
suction-flow quantities on the force measurements,
visual-flow studies, and surface-pressure-distribution
measurements at zero yaw over a range of angle of
attack. Reynolds number effects were determined
from 3.0 x 106 to 7.5 x 106 for the wing in the sealed
and faired condition.


NACA RM L50CO2a

HIGH-SUBSONIC PERFORMANCE CHARACTERIS-
TICS AND BOUNDARY-LAYER INVESTIGATIONS
OF A 120 10-INCH-INLET-DIAMETER CONICAL
DIFFUSER. B. H. Little, Jr. and Stafford W. Wil-
bur. May 11, 1950. 62p. diagrs., photos. (NACA
RM L50CO2a) (Declassified from Restricted,
6/11/53)




NACA
RESEARCH ABSTRACTS NO.46 |

An investigation was made of the effect of inlet-
boundary' -layer thickness a.id Mach number on the
performance of a 12' 10-inch-inlet-diameter conical
diffuser of 2.1 area ratio. Inlet Mach number was
varied from 0. 2 to choking and Reynolds number
(based on inlet diameter) from 1 x 106 to 3.9 x 106.
Performance coefficients, static-pressure distribu-
tions, boundary-layer profiles at seven stations in
the diffuser, and boundary-layer parameters are
presented

NACA RM L52G31

LALrLiBHATION OF STRAIN-GAGE INSTALLATIONS
IN AIRCRAFT STRUCTURES FOR THE MEASURE-
MENT OF FLIGHT LOADS. T. H. Skopinski,
William S. Aiken, Jr. and Wilber B. Huston.
October 8, 1952. 71p. diagrs., 10 tabs. (NACA
RM L52G31) (Declassified from Confidential,
6/29/53)

A general procedure is developed for calibrating
strain-gage installations in aircraft structures for
application to flight measurements of loads. The
basic procedure can be modified as necessary to suit
the requirements of any particular structure. The
application of the procedure is illustrated by results
for two typical structures.


NACA-Laugley 7-31-53 4D00





UNIVERSITY OF FLORIDA


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