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National Advisory Committee for Aeronautics
JUNE 27, 1955
CURRENT NACA REPORTS
NACA Rept. 1187
THEORETICAL AND EXPERIMENTAL INVESTIGA-
TION OF ADDITIVE DRAG. Merwin Sibulkin. 1954.
ii, 12p. diagrs. (NACA Rept. 1187. Formerly
The significance of additive drag is discussed and
equations for determining its approximate value are
derived. Charts are presented giving values of ad-
ditive drag for open-nose inlets and for annular-nose
inlets with conical flow at the inlet. The effects 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-
NACA RM L55E09a
SUMMARY OF RECENT THEORETICAL AND
EXPERIMENTAL WORK ON BOX-BEAM VIBRA-
TIONS. JohnM. Hedgepeth. June 1955. 10p.
diagrs., photo., 2 tabs. (NACA RM L55E09a)
A discussion of various secondary effects which have
an important influence on the vibration characteris-
tics of box beams is presented. Means of incorporate -
ing these effects in vibration analyses of actual built-
up box beams are discussed. Comparisons with
experiment are given; good agreement between
theory and experiment is obtained when the secondary
effects are included.
---. "C'A"RM L55Ellb
PRELIMINARY INVESTIGATION OF THE COMPRES-
SIVE STRENGTH AND CREEP LIFETIME OF
S2024-T3 (FORMERLY 24S-T3) ALUMINUM-ALLOY
PLATES AT ELEVATED TEMPERATURES. Eldon
E. Mathauser and William D. Deveikis. June 1955.
12p. diagrs. (NACA RM L55Ellb)
The results of elevated-temperature compressive
strength and creep tests of 2024-T3 (formerly
245-T3) aluminum-alloy plates supported in V-
grooves are presented. A relation previously de-
veloped for predicting plate compressive strength at
room temperature was satisfactory for determining
elevated temperature strength. Creep-lifetime re-
sults are presented for the plates in the form of
master creep-lifetime curves using a time tempera-
ture parameter. A comparison is made between
tensile and compressive creep lifetime for theplates,
and the magnitude by which the design stress is
decreased because of material creep and Loss ot
strength due to exposure at elevated temperatures is
NACA RM L55E12b
ALS UNDER RA
George J. Hein
0lp. diagrs. (
TRTIES OF SOME SHEET MATERI-
ierl ind John E. Inge. June 1955.
4ACA RM L55E12b)
A ".-, .. .- '
Results are prdsentpd of-tests to-deterufine e ef
fect of heating it un for ro
0.20 F to 1000 Y per s n~ roper s
of some sheet aaterAKIUn 'cfnSfjr rnnd
lions 7075-T6M 75=-T6) and 2024-T3 (24S-T3)
aluminum alloys, Inconel, and RS-120 titanium al-
loy. Some comparisons are given between yield and
rupture stresses obtained under rapid-heating con-
ditions and those obtained from elevated-temperature
stress-strain tests for 1'2-hour exposure. Master
yield- and rupture-stress curves based on the use of
a linear temperature-rate parameter are included
which provide a convenient method for predicting
yield and rupture stresses and temperatures for'dif-
ferent temperature rates.
NACA TM 1341
APPROXIMATE HYDRODYNAMIC DESIGN OF A
FINITE SPAN HYDROFOIL. (Priblizhennyi
gidrodinamicheskii raschet podvodnogo kryla
konechnogo razmakha). A. N. Vladunirov. June
1955. 68p. diagrs., 5 tabs. (NACA TM 1341.
Trans. from Central Aero-Hydrodynamical Institute,
Rept. 311, 1937).
Previous work on the motion of various bodies under
the surface of a heavy fluid is di The solu-
tion of the motion of a flat pl by.-NdMI d %.-
Lavrentiev is applied to t L. o oi a ,
making possible the pre dia1pr on of charts
determining the lift and eWSitance of an infini6 an
hydrofoil operating in a I ion ss fluid ng
infinite depth below the a water sulacW5Son-
sideration is given to th eitcts of viscosity and4
method is suggested to drr t for the finite pa
The effect of the water s fac on the dow WasK
behind the foil is also disc ed.4 arit of
theoretical results obtained f this wor ith
experimental data indicates that a or the ap-
proximate hydrodynamic design of a finite span
hydrofoil has been achieved.
NACA TN 3378
ACOUSTICAL TREATMENT FOR THE NACA 8- BY
6-FOOT SUPERSONIC PROPULSION WIND TUNNEL.
Leo L. Beranek, Samuel Labate and Uno Ingard,
Bolt Beranek and Newman, Inc. June 1955. 86p.
diagrs., photo., 7 tabs. (NACA TN 3378)
"AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST, NW.,
; THE REPORT TITLE AND AUTHOR.
WASHINGTON 5s, D. C., CITING CODE NUMBER ABOVE EACH TITLE;
This report summarizes the results of a project at
the Lewis Flight Propulsion Laboratory to silence
the 8- by 6-foot supersonic wind tunnel. Sound
measurements in the neighborhood surrounding the
tunnel were conducted to evaluate the noise-attenuation
requirements. A muffler-development program was
continued until these attenuations were achieved.
The final design for the acoustic treatment is des-
cribed and experimental performance curves are
compared with anticipated theoretical results.
NACA TN 3382
EXPERIMENTS WITH A ROTATING-CYLINDER
VISCOMETER AT HIGH SHEAR RATES. J. A. Cole,
R. E. Petersen and H. W. Emmons, Harvard
University. June 1955. 31p. diagrs., tab. (NACA
Two straight mineral oils and a polymer-containing
oil have been tested in a rotating-cylinder viscome-
ter at high shear rates (maximum 0.25 million
reciprocal seconds) and the accompanying heat ef-
fects have been investigated. The torque measure-
ments are of low accuracy and fail to establish the
moderately non-Newtonian behavior of the polymer-
containing oil, but the temperature measurements,
which are in good agreement with a thermal analy-
sis, do indicate the presence of temporary viscosity
decrease at high shear rates for this oil.
NACA TN 3411
PRESSURE WAVES GENERATED BY ADDITION OF
HEAT IN A GASEOUS MEDIUM. Boa-Teh Chu,
Johns Hopkins University. June 1955. 47p. diagrs.
(NACA TN 3411)
The approximate formula of a linearized solution for
the pressure field generated by a moderate rate of
heat release is given. The analogies between the
pressure waves generated by heat release and those
generated by (1) mass release, (2) piston motion, or
(3) a two-dimensional body in a supersonic stream
are established analytically. The exact solution of
an idealized problem in which heat is released uni-
formly at a section of tube with a given rate, large
or small, is also constructed. The corresponding
problems in three dimensions are also solved. Some
applications of the theory are given.
NACA TN 3412
CREEP AND CREEP-RUPTURE CHARACTERISTICS
OF SOME RIVETED AND SPOT-WELDED LAP
JOINTS OF AIRCRAFT MATERIALS. Leonard
Mordfin, National Bureau of Standards. June 1955.
53p. diagrs., photos., 6 tabs. (NACA TN 3412)
Equipment, test techniques, and results are pre-
sented for an experimental investigation of the creep
of lap joints. Riveted aluminum-alloy joints fabri-
cated from 75S-T6 and 24S-T3 sheet with 24S and
24S-T31 rivets were tested at 3000, 4000, and 5000
F. Spot-welded joints of 1/4-hard, type 301 stain-
less steel were tested at 8000 F. Each type of joint
was also tested in tension at room temperature.
RESEARCH ABSTRACTS NO.85
NACA TN 3416
THEORETICAL AND EXPERIMENTAL INVESTIGA-
TION OF THE EFFECT OF TUNNEL WALLS ON
THE FORCES ON AN OSCILLATING AIRFOIL IN
TWO-DIMENSIONAL SUBSONIC COMPRESSIBLE
FLOW. Harry L. Runyan, Donald S. Woolston and
A. Gerald Rainey. June 1955. 41p. diagrs., photo.
(NACA TN 3416)
The integral equation defining the problem of an
oscillating wing in a tunnel is treated and is pre-
sented in a form adapted to calculations. Applica-
tion is made to a number of examples to illustrate
the influence on the magnitude of wall effects of
variations in frequency, Mach number, and ratio of
tunnel height to wing semichord. Comparison is
made with experimental measurements for several
subsonic Mach numbers.
NACA TN 3448
THEORETICAL ANALYSIS OF INCOMPRESSIBLE
FLOW THROUGH A RADIAL-INLET CENTRIFUGAL
IMPELLER AT VARIOUS WEIGHT FLOWS. I -
SOLUTION BY A MATRIX METHOD AND COMPAR-
ISON WITH AN APPROXIMATE METHOD. Vasily D.
Prian, James J. Kramer and Chung-Hua Wu. June
1955. 39p. diagrs., tab. (NACA TN 3448)
A method for the solution of the incompressible, non-
viscous flow through a centrifugal impeller, including
the inlet region, is presented. Several numerical
solutions are obtained for four weight flows through
an impeller at one operating speed. The results are
presented in a series of figures showing streamlines
and resultant velocity contours. A comparison is
made with the results obtained by use of a rapid ap-
proximate method of analysis.
NACA TN 3449
THEORETICAL ANALYSIS OF INCOMPRESSIBLE
FLOW THROUGH A RADIAL-INLET CENTRIFUGAL
IMPELLER AT VARIOUS WEIGHT FLOWS. H -
SOLUTION IN LEADING-EDGE REGION BY RELAX-
ATION METHODS. James J. Kramer. June 1955.
19p. diagrs. (NACA TN 3449)
The detailed solution of the flow around the blade
nose of a 48-inch-diameter radial-inlet centrifugal
impeller has been obtained by relaxation methods for
four weight flows. The results are presented in a
series of figures showing streamlines and relative
velocity contours. Minimum velocity gradients
around the blade nose occurred for the weight flow
corresponding to a mean angle of attack of -4.60
computed from blade speed and an upstream axial-
radial velocity for which blade blockage has been
taken into account. A small positive local angle of
attack seems desirable for blades with rounded
NACA TN 3457
ESTIMATION OF INLET LIP FORCES AT SUBSONIC
AND SUPERSONIC SPEEDS. W. E. Moeckel. June
1955. 12p. diagrs. (NACA TN 3457)
RESEARCH ABSTRACTS NO. 85
The effects of inlet lip thickness on inlet performance
are estimated as functions of mass flow for subsonic
and supersonic flight speeds. At subsonic speeds,
pressure-recovery losses and additive drag are
shown to decrease linearly with increasing lip frontal
area if the maximum suction force is attained. Al
supersonic speeds, inlet drag increases linearly
with inlet lip frontal area at iull mass flow. For
reduced mass flow, some reduction in additive drag
is possible with lips of moderate thickness, but the
magnitude of this reduction becomes negligible as
flight speed increases.
NACA TN 3461
AT A MACH NUMBER OF 1.62. Maurice J.
Brevoort and Bernard Rashis. June 1955. 15p.
diagrs., tab. (NACA TN 3461)
An axially symmetric annular nozzle was used to
obtain essentially flat-plate results for turbulent-
heat-transfer coefficients and temperature-recovery
factors. The test results are for a Mach number of
1.62 and for a Reynolds number range of 7.22 x 105
to 1.20 x 108. The heat-transfer-coefIicient results
agree with theoretical results for M = 1.60 and
T,'/T = 1.60. The recovery factors are on the
average 1.5 percent lower than data for a Mach num-
ber of 2.4.
NACA TN 3499
CALCULATION OF THE SUPERSONIC PRESSURE
DISTRIBUTION ON A SINGLE-CURVED TAPERED
WING IN REGIONS NOT INFLUENCED BY THE ROOT
OR TIP. Walter G. Vincenti and Newman H. Fisher,
Jr. June 1955. 32p. diagrs. (NACA TN 3499)
The shock-expansion method for the calculation of
the pressures on cylindrical wings in supersonic
flow is extended to tapered wings made up of single-
curved surfaces. The method applies in regions
where (a) the component of velocity normal to the
surface rulings is supersonic and (b) the flow is not
influenced by tips or junctures. In these regions the
'flow is defined by a pair of ordinary differential
equations whose solution is readily obtained by nu-
merical means. Results are shown and discussed
for a representative triangular wing.
NACA TN 3502
THE TRANSONIC CHARACTERISTICS OF 38
CAMBERED RECTANGULAR WINGS OF VARYING
ASPECT RATIO AND THICKNESS AS DETERMINED
BY THE TRANSONIC-BUMP TECHNIQUE. Warren
H. Nelson and Walter J. Krumm. June 1955. 173p.
diagrs., photos. (NACA TN 3502. Formerly
An investigation was made in the Ames 16-foot high-
speed wind tunnel utilizing the transonic-bump tech-
nique to determine the aerodynamic characteristics
at transonic Mach numbers of 38 cambered rectangu-
lar wings. The wings had aspect ratios of 4, 3, 2,
1.5, and 1, and NACA 63A2XX and 63A4XX sections
with thickness-to-chord ratios of 10, 8, 6, 4, and 2
percent. The Mach number range was 0.6 to 1.12
with corresponding Reynolds numbers of 1.7 to 2.2
million. The data are presented without analysis.
Aeroplane and Armament Experimental Establish-
ment (Gt. Bril.) DETERMINATION OF THE VARI-
ATION WITH ALTITUDE OF THE RECOVERY
FACTOR OF AIR THERMOMETERS. December 17,
1954. 28p. diagrs., photos., 10 tabs. (AAEE
Tests were made for the standard knife edge and two
other available air thermometer bulbs. The results
show a significant decrement with altitude in the
value of the recovery factor of the order of 0.2 be-
tween 5.000 and 40,000 feet. That means a change
from 0.69 to 0.51 in the values for the standard knile
edge bulb. In view of these findings, it appears
imperative in performance tests which are sensitive
to ambient air temperature, to determine the recov-
ery factor of the outside air thermometer over the
altitude range of the aircraft. Alternatively, there
is a need for an air thermometer whose recovery
factor remains constant with change of altitude.
Royal Aircraft Establishment (Gt.Brit.)
THE TORSIONAL VIBRATIONS OF A CLASS OF
THIN, TAPERED, SOLID WINGS. Elizabeth A.
Frost. January 1955. 22p. diagrs., tab. (RAE
Tech. Note Structures 152)
This report considers the torsional vibrations of
thin solid wings of doubly symmetrical chordwise
section, with linear variation of chord and parabolic
variation of thickness. Frequencies of symmetrical
and antisymmetrical vibrations are presented
graphically for a range of values of the aspect ratio
and the taper ratio.
Royal Aircraft Establishment (Gt. Brit.)
TESTS ON THE PROTECTION GIVEN TO METALS
BY ETCH PRIMERS. H. G. Cole. January 1955.
13p. 6 tabs. (RAE Tech. Note Met. 209)
Long term seawater spray corrosion tests have
shown that the relative performance of painting
schemes based on six commercial etch primers
varied markedly from metal to metal. On zinc and
cadmium-plated steel,the etch primer schemes gave
better performance than conventional paints applied
over classical methods of surface treatment. On
steel and aluminum alloys, the use of etch primers
is an acceptable substitute for conventional methods
of protection, provided that the etch primer is fol-
lowed by a full protective scheme. Their use on
magnesium alloys is risky because of the danger of
attack on the metal.
Royal Aircraft Establishment (Gt. Brit.)
A JUNCTION TRANSISTOR DECADE COUNTER.
H. W. P. Knapp. January 1955. lip. diagrs.,
photo. (RAE Tech. Note GW 354)
A junction transistor decade counter is described
which is tolerant of supply voltage, input amplitude,
and ambient temperature variations.
DECLASSIFIED NACA REPORTS
NACA RM A52G17
THE USE OF LEADING-EDGE AREA SUCTION TO
INCREASE THE MAXIMUM LIFT COEFFICIENT OF
A 350 SWEPT-BACK WING. Curt A. Holzhauser and
Robert K. Martin. September 1952. 37p. diagrs.,
photo., 3 tabs. (NACA RM A52G17) (Declassified
from Confidential, 6/10/55)
Measurements were made of the increase in lift co-
efficient obtained with various amounts of area suc-
tion. Flow coefficients and power inputs required to
obtain these lift coefficients were measured at free-
stream velocities from 112 to 180 feet per second.
With full-span area suction and a maximum chord-
wise opening of 2.2 percent, the CLmax of the model
with flaps deflected was increased from 1.33 to 2.00
with a flow coefficient of 0.00108 and a suction horse-
power of 47 at a free-stream velocity of 129 feet per
second. There appeared to be no unacceptable char-
acteristics that a 350 swept-wing airplane would
exhibit in low-speed flight as a result of employing
NACA RM L7F30
FREE-FLIGHT INVESTIGATION OF CONTROL EF-
FECTIVENESS OF FULL-SPAN, 0.2-CHORDPLAIN
AILERONS AT HIGH SUBSONIC, TRANSONIC, AND
SUPERSONIC SPEEDS TO DETERMINE SOME EF-
FECTS OF WING SWEEPBACK, TAPER, ASPECT
RATIO, AND SECTION-THICKNESS RATIO. Carl A.
Sandahl. August 13, 1947. 16p. diagrs., photos.,
tab. (NACA RM L7F30) (Declassified from
The aileron control characteristics of untapered, 450
sweptback wings of aspect ratio 3 were found to be
generally the same for the 65-006 and 65-009 sec-
tions. The tapered 450 sweptback wings of aspect
ratio 3 and 65-009 section exhibited a small abrupt
change in rolling effectiveness from M = 0.92 to 1.00
which was not characteristic of untapered wings
tested having the same sweep, aspect ratio, and
section. A reduction in aspect ratio from 3 to 1.75
for unswept, untapered wings gave an increase in
rolling effectiveness. Both aspect-ratio configura-
tions showed undesirable control characteristics at
NACA RM L51F11
RECENT EXPERIMENTAL FLUTTER STUDIES.
Arthur A. Regier and Dennis J. Martin. June 12,
1951. 18p. diagrs. (NACA RM L51F11) (Declas-
sified from Confidential, 6/10/55)
This paper presents in a rather brief fashion some of
the highlights of recent experimental flutter studies.
Material from several papers is brought together
and compared. The material includes trend studies
of flutter of swept and unswept wings at transonic
speeds; studies of a bending-type flutter of swept
wings; flutter involving body modes; and preliminary
studies of M and W and delta wing configurations.
RESEARCH ABSTRACTS NO. 85
NACA RM L51L03
PRESSURE DISTRIBUTIONS AT MACH NUMBERS
FROM 0.6'TO 1.9 MEASURED IN FREE FLIGHT ON
A PARABOLIC BODY OF REVOLUTION WITH
SHARPLY CONVERGENT AFTERBODY. William E.
Stoney, Jr. April 1952. 34p. diagrs., photos.
(NACA RM L51L03) (Declassified from Confidential,
A flight test was made, at Mach numbers from 0.6 to
1.9, of a fin-stabilized parabolic body of revolution
having a sharply convergent afterbody. Pressures
were measured at eight longitudinal stations on the
body and on the base with and without a simulated
wind-tunnel sting extending from the base of the
model. The results were compared with theoretical
determinations of the pressures by linear and exact
NACA RM L52L05
WIND-TUNNEL INVESTIGATION OF STALL CON-
TROL BY SUCTION THROUGH A POROUS LEADING
EDGE ON A 370 SWEPTBACK WING OF ASPECT
RATIO 6 AT REYNOLDS NUMBERS FROM 2.50x106
to 8.10 x 106. Robert R. Graham and William A.
Jacques. March 1953. 67p. diagrs.. photo., 2 tabs.
(NACA RM L52L05) (Declassified from Confidential,
Stall control by suction through a porous leading-
edge upper surface was investigated on a plain wing
and on the same wing with half-span split or double
slotted flaps. Some effects of varying the chordwise
and spanwise extent of porosity and varying the flow
coefficient were investigated on the plain wing.
Force and moment data are presented along with
minimum measured leading-edge pressure coeffi-
cients across the span of the wing.
NACA RM L53C02
THE BASE PRESSURE AT SUPERSONIC SPEEDS
ON TWO-DIMENSIONAL AIRFOILS AND BODIES OF
REVOLUTION (WITH AND WITHOUT FINS) HAVING
TURBULENT BOUNDARY LAYERS. Eugene S. Love.
April 1953. 65p. diagrs., photos.- (NACA
RM L53C02) (Declassified from Confidential,
An analysis is made of available experimental data
to show the effect of most of the variables that are
more predominant in determining base pressure at
supersonic speeds. The analysis is restricted to
turbulent boundary layers and covers two-
dimensional bases and the bases of bodies of revolu-
tion, with and without stabilizing fins. An analogy to
the pressure rise required to separate the boundary
layer is presented as are simple semiempirical
methods for the estimation of base pressure.
NACA RM L53F15a
STUDIES OF THE SPEED STABILITY OF A TAN-
DEM HELICOPTER IN FORWARD FLIGHT. Robert
J. Tapscott and Kenneth B. Amer. August 1953.
35p. diagrs., photos., tab. (NACA RM L53F15a)
(Declassified from Confidential, 6/10, 55)
RESEARCH ABSTRACTS NO. 85
Analytical and experimental studies and correspond-
ing pilots' opinions of the speed stability of a tandem
helicopter are presented and means for improving
the speed stability are discussed. An analytical ex-
pression is derived for use in predicting changes in
speed stability of a tandem helicopter brought about
by changes in the rotor geometry. Constants for use
with the analytical expression are presented in chart
NACA RM L53GI5a
THE VARIATION OF ATMOSPHERIC TURBULENCE
WITH ALTITUDE AND ITS EFFECT ON AIRPLANE
GUST LOADS. Robert L. McDougal, Thomas L.
Coleman and Philip L. Smith. November 1953. 16p.
diagrs., 2 tabs. (NACA RM L53G15a) (Declassified
from Confidential, 6/10 '55)
Analysis of turbulence data for altitudes up to 60,000
feet indicates substantial reductions in the amount
and relative mtensity of the turbulence at the higher
altitudes. The implications of these reductions in
the turbulence are discussed in regard to the gust
loads for assumed airplane operations.
NACA RM L54B16b
FLIGHT TESTS OF A 0.4-SCALE MODEL OF A
STAND-ON TYPE OF VERTICALLY RISING AIR-
CRAFT. Marion O. McKinney and Lysle P. Parlett.
March 1954. 23p. diagrs., photos. (NACA
RM L54B16b) (Declassified from Confidential,
The results of a free-flight investigation of the sta-
bility and control of a 0.4-scale model of a stand-on
type of vertically rising aircraft in take-offs and
landings and m hovering and forward flight are pre-
sented. The aircraft component of the model con-
sisted of a motor-driven, single-rotation propeller
in a short shroud with antitorque vanes and control
surfaces at the rear of the shroud.
NACA-Langley 6-27-55 4M
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