This item is only available as the following downloads:
.National Advisory Committee for Aeronautics
NOVEMBER 9, 1954
CURRENT NACA REPORTS
NACA Repl. 1145
A METHOD OF CALIBRATING AIRSPEED INSTAL-
LATIONS ON AIRPLANES AT TRANSONIC AND
SUPERSONIC SPEEDS BY THE USE OF ACCEL-
EROMETER AND ATTITUDE-ANGLE MEASURE-
MENTS. John A. Zalovcik, Lindsay J. Lina and
James P. Trant, Jr. 1953. ii, 13p. diagrs.,
photos., tab. (NACA Rept. 1145. Formerly
TN 2099; TN 2570)
A method is described for calibrating airspeed
installations on airplanes at transoic and supersonic
speeds in which use is made of norflai and longitu-
dinal accelerations and altitude angle as measuredby-
instruments carried within the airplane. a rpp4e-li .
calibration of the pitot-static installati iokr a M
fighter airplane is presented as an experimental
check on the accuracy of the method.
NACA Rept. 1160
THE ZERO-LIFT DRAG OF A SLENDER BODY OF
REVOLUTION (NACA RM-10 RESEARCH MODEL)
AS DETERMINED FROM TESTS IN SEVERAL WIND
TUNNELS AND IN FLIGHT AT SUPERSONIC
SPEEDS. Albert J. Evans. 1954. ii, 13p.
diagrs., tab. (NACA Rept. 1160. Formiperly
Presents zero-Lit drag data of an NACA RM-10
slender body of revolution with and without stabi -
lizing fins attached. The results from several wind
tunnels and in flight are compared. The results
cover a Reynolds number range from about I x 106
to 40 x 106 for the wind-tunnel models and 12 x 10
to 140 x 106 for the flight models. The Mach
numbers covered include 1.5 to 2.4 in the wind
tunnels and 0.85 to 2.5 in flight.
NACA RM 54F22
EFFECTS OF MULTIAXIAL STRETCHING ON CRAZ-
ING AND OTHER PROPERTIES OF TRANSPARENT
PLASTICS. Irvin Wolock and Desmond A. George,
National Bureau of Standards. October 1954. 34p.
diagrs., photos., 10 tabs. (NACA RM 54F22)
An investigation was made of the effects of onrienta-
tion by multiaxial stretching on properties of various
plastic glazing materials. The materials studied
were Lucite HC-222 polymethyll methacrylate),
Plexiglas 55 (modified polymethyl methacrylate),
Gafite, and resin C polymethyll alpha-chloroacrylate).
The following tests were conducted on samples of
these materials stretched up to 150 percent: Di-
mensional stability at elevated temperatures, surface
abrasion, standard tensile tests, and stress-solvent
crazing tests using ethylene dichloride.
*AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST, NW.,
THE REPORT TITLE AND AUTHOR.
S2f /J 2.
NACA RM E54H04
SLOWING-DOWN DISTRIBUTION TO [NDIUM
RESONANCE OF NEUTRONS FROM A Ra-a-Be
SOURCE IN WATER-IRON MIXTURES. Daniel
Fieno. November 1954. 16p. diagrs., photo.,
2 tabs. (NACA RM E54H04)
The mean s uare slowing-down distance to indium
resonance r2 derived from the slowing-down
distribution has been measured for water and for
three water-iron mixtures for neutrons from a 0.1-
gram Ra-a-Be source. The values of r2 for water-
iron volume ratios of 1. 2, and 3 and for water were
347, 336, 314. and 291 centimeters squared, re-
spectively. Within the accuracy of the measurements,
the relaxation length X was approximately the same
for water and for the three water-iron mixtures, the
average value being 10.0 centimeters.
NACA TN 3152
TRANSVERSE OSCILLATIONS IN A CYLINDRICAL
COMBUSTION CHAMBER. Franklin K. Moore and
Stephen H. Maslen. October 1954. 25p. diagrs.
(NACA TN 3152)
Transverse oscillations in combustion chambers are
studied. With an axnal temperature gradient con-
sidered, the modes of weak oscillation are described
for low Mach numDer. Amplification by coupLing
with vigorous combustion in the flame-holder wake is
assumed. The amplification depends on flame-holder
and centerbody diameters and on time-lag effects.
The nature of finite transverse periodic waves is
also analyzed. Results show that such waves have
frequencies independent of amplitude and do not
steepen with time.
NACA TN 3302
LIQUEFACTION OF AIR IN THE LANGLEY 11-INCH
HYPERSONIC TUNNEL. Charles H. McLellan and
Thomas W. Williams. October 1954. 36p. diagrs.,
4 Labs. (NACA TN 3302)
Pressure and scattered-light measurements were
made in the Langley 1 1-inch hypersonic tunnel to
determine the effect of stagnation temperature on the
flow in two Mach number 7 nozzles and to determine
the nature of the condensation process occurring at
low stagnation temperatures. Liquefaction of the
air occurred very close to the saturation point with-
out a condensation shock. This result indicates that
liquefaction took place on foreign nuclei such as
water and carbon-dioxide particles. The results
from varying the water vapor and carbon-dioxide
content, however, could not be correlated with Max
Volmer's condensation theory. The average particle
radius was 480 angstroms in the test section of the
WASHINGTON 25 D C, CITING CODE NUMBER ABOVE EACH TITLE,
me conditions about lul'-' p.a rn ic i. per
ter were present.
TAL DETERMINATION OF BOUNDARY-
'SirlON ON A BODY OF REVOLUTION
James R. Jedlicka, Max E. Wilkins
t. October 1954. 56p. diagrs., photos.
42. Formerly RM .; L1 |
er-cransition tests were made in free
i still air on a small-scale fin-stabilized
,tion of fineness ratio 30 at a Mach
and length Reynolds numbers of 12 and
bhree types of surfaces were tested and
Reynolds number was found to depend
he surface smoothness. Angle of attack
'fluence transition strongly. This
related in terms of pressure rise along
nd led to a prediction of the effect of
dy shape on sheltered-side transition.
Transition points were time dependent,
one case over a range of Reynolds
S4 io 20 million. Steady laminar flow
o a. length Reynolds number of 11
isition due to roughness and adverse
ient occurred even though the tests
- region of theoretical infinite laminar
Research Council (Gt.Brit.)
'EED LABORATORY OF THE AERO-
!VISION, N.P.L. D. W. Holder.
des rs., photos,, 2 tabs. (ARC
ec tunnel installation of the Aerodynam-
Sthte National Physical Laboratory is
hlie installation consists of the 12-in.
-sceed tunnel, the 20- by 8-in. high-
ad a number of smaller tunnels all of
rated on the induction principle from a
i-essed-air storage capacity. The new
nagh-speed tunnel is also described.
%perimental techniques which have been
1, including the schlieren and
i:ethods. The last of the report re-
prin-ental results obtained in the high-
.cring and immediately before the war.
a which occur on a particular airfoil
reaised are described and the effect of
Research Council (Gt. Brit.)
CREASE AT HIGH SUBSONIC SPEEDS.
1954. 16p. diagrs. (ARC R & M
246. Formerly RAE Tech. Note
RESEARCH ABSTRACTS NO.73
The drag increase beyond the critical Mach number
is calculated by modifying the supersonic part of the
Karman-Tsien pressure distribution on a profile.
This is possible when the supersonic regions are not
too large. The formula giving the modified pressure
distribution is derived very roughly. It may give
only one of the main effects appearing when super-
sonic speeds occur in the flow, and may be changed
and calculated more exactly later. For the calcula-
tion of the drag-increase, the formula is sufficient.
Within the approximation of the theory the lift coef-
ficient is practically unchanged.
Aeronautical Research Council (Gt. Brit.)
A SIMPLE METHOD OF COMPUTING CD FROM
WAKE TRAVERSES AT HIGH-SUBSONIC SPEEDS.
J. S. Thompson. 1954. 12p. diagrs., 4 tabs.
(ARC R & M 2914; ARC 8462. Formerly RAE
This note gives a convenient method of obtaining CD
from a pitot-static traverse in an airfoil wake, using
Jones' modified equation for compressible flow.
Charts are provided from which the integrand CD'
can easily be obtained for any point in the traverse,
but it is shown that in nearly all cases an accuracy
of 1 percent in CD can be obtained by applying an
integrating factor to the area under the total-head
loss curve. Three appendices give (a) a summary of
the standard theory and equations, (b) details of the
construction of the charts,and (c) an empirical
equation giving CD'in a simple analytical form.
Aeronautical Research C in- (Gt. Brit.)
NOTE ON THE DYNAMIC CHARACTERISTICS OF
SERVO-TAB SYSTEMS OF CONTROL. D. Adamson
and D. J. Lyons. 1954. 12p. diagrs., tab. (ARC
R& M 2853; ARC 11,666. Formerly RAE Aero 2263)
Curves have been constructed from which estimates
can be made of those dynamic characteristics of the
servo-tab-type of control which are of chief interest
to 'r .dsLi'icner, namely, the magnitude of the first
overshoot of the main flying control beyond its equil-
ibrium *-sir i..n. the lag of the main control surface
behind the tab movement, the damping of the main
control surface oscillation, and the angular velocity
possessed by the nain control when it first passes
liI.I.I i its equilibrium position.
Aeronautical Resear h Council (Gt. Brit.)
MODEL TESTS OF AN AIR INTERCHANGE SYSTEM
FOR REMOVING ENGINE EXHAUST PRODUCTS
FROM A WIND TUNNEL. K. W. Newby, E. G.
Barnes and D. W. Bottle. 1954. 32p. diagrs.,
3 tabs. (ARC R M 2639; ARC 11,604. Formerly
RAE Aero 2249)
An invi,,t .,Inii- has been made on the lun LI j.nI1g of
an air interchange system for removing from a
return-circuit wind tunnel a high proportion of the
exhaust products from propulsive units under test.
RESEARCH ABSTRACTS NO. 73
The tests were planned to assist the design of an
engine altitude tunnel. The system tested was gen-
erally very satisfactory for the specified require-
ments, and can operate up to interchange ratios of
the order of 15 percent without Interfering appreci-
ably with the flow in the tunnel working section.
ON BRANCHED POTENTIALS IN SPACE. (tber
verzweigte Potentiale in Raum). A. Sommerfeld.
September 1954. 36p. diagrs. (Trans. from London
Mathematical Soc., Proceedings, v. 28, April 8,
Thomson's method of images and its extension by
means of branched potentials is discussed. Green's
function of a Riemann space with a simple recti-
linear branch curve; applications of Green's function
of the winding space to problems of ordinary
potential theory; and Green's function of a Riemann
space with two parallel rectilinear branching curves
and its applications are also discussed.
ON THE STRESS ANALYSIS OF SWEPT WINGS.
(Over de sterkteberekening van pijIvormige vleugels).
J. P. Benthem. September 1954. 182p. diagrs.
(Trans. from Nationaal Luchtvaartlaboratorlum,
Amsterdam. S. 405)
Methods of stress analysis of swept wings are
discussed and reviewed. Many of the references
are commented upon.
DECLASSIFIED NACA REPORTS
THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL, 10/12/54:
NACA RM A7A15
FLIGHT-TEST MEASUREMENTS OF AILERON CON-
TROL SURFACE BEHAVIOUR AT SUPERCRITICAL
MACH NUMBERS. Harvey H. Brown, George A.
Rather, Jr. and Lawrence A. Closing. April 23,
1947. 26p. diagrs., photos., 2 tabs. (NACA
RM A7A15 (Declassified from Confidential,
The behavior at supercritical Mach numbers of the
ailerons of a jet-propelled fighter has been measured
.1p to 0.866 Mach number. The considerable amount
oi aileron upfloat occurring at these Mach numbers
was found to be due to a large loss in pressure re-
covery on the upper surface aft of the shock wave
which caused very large increases in the aileron
hinge moments. Data obtained from pressure distri-
butiion measurements are presented to show the very
critical effect of Mach number on the magnitude of
thesee hinge moments.
NACA RM A7G03
AN ANALYSIS OF LONGITUDINAL-COT
PROBLEMS ENCOUNTERED IN FLIGH'
SONIC SPEEDS WITH A JET-PROPELL
PLANE. Harvey H. Brown, L. Stewart
Lawrence A. C'l 'i-. i September 25,
diagrs.. photos., 3 tabs. (NACA RM A
(Declassified from Confidential, 10/ i2
During flight tests of a jet-propelled ai
den pitch- -up motion of the airplane oec
recovery from a high-speed dive, althoi
had not moved the controls so as to proi
motion. Measurements of the stability
characteristics of the airplane and of th
sure distribution during the dive and r.k
NACA RM A7I16
HIGH-SPEED WIND-TUNNEL TESTS 0
PURSUIT AIRPLANE AND CORRELATE
FLIGHT-TEST RESULTS. Joseph W.
Lyle J. Gray. January 21, 1948. 56p
photos. (NACA RM A711 i; (Declassif
A wind-tunnel investigation of the aero
characteristics of a 1 3- a le model r
pelled airplane was made at high subsi
bers for comparison with flight test re
were made of wing and fuselage dive-r
to determine their effectiveness at higl.
general, the wind-tunnel results show
ment with the flight-test data. Both wi
lage dive-recovery flaps were effective
ing longitudinal control. Huw,.ver. the
recovery flaps lost their effectiveness
NACA RM A51E04
EXPERIMENTAL STUDY OF THE EFI
SWEEPBACK ON TRANSONIC AILERO
Lionel L. Levy, Jr. and Earl D. Knelct
September 1951. 20p. diagra, photo.
A51E04) (Declassified from Confidenti
The effect of sweepback on the Mach no
which transonic aileron flutter occurs
mined for a wing -aileron combination
65-213, a = 0.5, airfoil section. The
range of the investigation extended fro
approximately and the Reynolds number
million. Angle of attack was varied fr
for sweep angles of 00 to 500.
NACA RM A51G10
PRELIMINARY INVESTIGATION OF T1
FLUCTUATIONS IN THE WAKES OF T
DIMENSIONAL WINGS AT LOW ANGEL
TACK. Robert M. Surenson, John A
James C. Kyle. October 1951. 58p.
photos. (NACA RM A51G10). (Declas&
An experimental investigation has beet
obtain fundamental two-dimensional da
tent, magnitude, and other character
pressure fluctuations behind airfoils with NACA
23013 and NACA 651-213 sections. Results are pre-
sented for the airfoil with the NACA 23013 section
through a range of Mach numbers from 0.60 to 0.80
and through a range of angles of attack of -20 to +50.
The report discusses methods of defining wing- and
tail-buffet boundaries in terms of pressure fluctua-
tions in the wake.
NACA RM A51H15
LOAD DISTRIBUTION OVER A FUSELAGE IN COM-
BINATION WITH A SWEPT WING AT SMALL
ANGLES OF ATTACK AND TRANSONIC SPEEDS.
Maurice D. White and Bonne C. Look. November
1951. 26p. diagrs., photo., tab. (NACA RM
A51H15) (Declassified from Confidential, 10/12/54)
Free-fall tests were made at Mach numbers from
0.85 to 1.06 of a wing-body configuration having a 450
sweptback cambered and twisted wing of aspect ratio
6 on a fuselage of fineness ratio 12.4. The distribu-
tions of load over the fuselage as determined from
surface pressure measurements are shown in rela-
tion to the lifts of the wing for small angles of attack.
Comparisons are made with wind-tunnel load distri-
butions and with theoretical load distributions for
subsonic and supersonic speeds.
NACA RM A51I25
AN ANALYSIS OF THE EFFECT OF STRUCTURAL
FEEDBACK ON THE FLUTTER OF A CONTROL
SURFACE HAVING POWER-BOOST SYSTEM.
Robert H. Barnes. June 1952. 29p. diagrs., photos.
(NACA RM A51I25) (Declassified from Confidential,
A wind-tunnel program was conducted to determine
the cause of flutter which had been experienced on an
airplane which employed a full power-boost system.
Analysis of the data showed that the power-boost
control valve was being actuated due to structural
deformation. Accordingly, an analytical study was
conducted which showed that structural feedback
could have been the cause of flutter.
NACA RM A51K27
BODIES OF REVOLUTION FOR MINIMUM DRAG AT
HIGH SUPERSONIC AIRSPEEDS. A. J. Eggers, Jr.,
David H. Dennis and Meyer M. Resnikoff. February
1952. 44p. diagrs., photos. (NACA RM A51K27)
(Declassified from Confidential, 10/12/54)
Newtonian impact theory is used to determine body
shapes of minimum drag under various combinations
of the conditions of given body length, base diameter,
surface area, and volume. In addition an estimate
is made of centrifugal forces, and their effect on one
minimum-drag shape is considered. An experimen-
tal investigation carried out in the Ames 10- by 14-
inch supersonic wind tunnel on a family of bodies,
including two of the minimum-drag shapes, is found
to provide a substantiation of the analysis.
RESEARCH ABSTRACTS NO. 73
NACA RM A51LO3a
SOME EFFECTS OF SIDE-WALL MODIFICATIONS
ON THE DRAG AND PRESSURE RECOVERY OF AN
NACA SUBMERGED INLET AT TRANSONIC SPEEDS.
Robert A. Taylor. February 1952. 25p. diagrs.,
photos. (NACA RM A51LO03a. (Declassified from
Comparative data were obtained for an NACA sub-
merged inlet and two ramp-wall modifications of the
NACA submerged inlet. These modified inlets were
generally superior to the NACA inlet from the stand-
point of pressure recovery at the highest test mass-
flow ratios. No significant changes in drag were
produced by the modification for Mach numbers be-
low 1.0 and for a mass-flow ratio of 0.88, but small
increases in drag, at supersonic Mach numbers and
the higher angles of attack, resulted from the modi-
NACA RM A51L17b
A CORRELATION BY MEANS OF THE TRANSONIC
SIMILARITY RULES OF THE EXPERIMENTALLY
DETERMINED CHARACTERISTICS OF 22 RECTAN-
GULAR WINGS OF SYMMETRICAL PROFILE.
John B. McDevitt. February 1952. 60p. diagrs.,
3 tabs. (NACA RM A51L17b) (Declassified from
The similarity rules have been used to correlate the
experimental data for a series of 22 rectangular,
symmetrical wings having NACA 63AOXX sections,
aspect ratios from 1/2 to 6, and thicknesses from 2
to 10 percent. The data were obtained by use of the
transonic bump technique over a Mach number range
of 0.40 to 1.10, corresponding to a Reynolds number
range from 1.25 to 2.05 million.
NACA RM A52A29
FLIGHT TESTING BY RADIO REMOTE CONTROL -
FLIGHT EVALUATION OF A BEEP-CONTROL
SYSTEM. Howard L. Turner, John S. White and
Rudolph D. Van Dyke, Jr. April 1952. 55p. diagrs,
photos., tab. (NACA RM A52A29) (Declassified
from Confidential, 10/12/54)
A comparison between manual control and remote
control showed that a beep-type, ratio-remote-
control system was, in general; a satisfactory means
of control for conducting standard handling-quality
flight tests. The dynamic characteristics of the
airplane-autopilot combination and the selection of
the proper parameter adjustments are discussed.
NACA RM A52B05
FULL-SCALE WIND-TUNNEL INVESTIGATION OF
THE EFFECTS OF WING MODIFICATIONS AND
HORIZONTAL-TAIL LOCATION ON THE LOW-
SPEED STATIC LONGITUDINAL CHARACTERISTICS
OF A 350 SWEPT-WING AIRPLANE. Ralph L. Maid.
April 1952. 54p. diagrs., photos., 7 tabs. (NACA
RM A52B05) (Declassified from Confidential,
RESEARCH ABSTRACTS NO. 73
Low-speed lift, drag, and pitching-moment charac-
teristics of a full-scale 350 swept-wing airplane are
presented for Reynolds numbers ranging from
3.2 x 106 to 12.3 x 106, and with the horizontal tail
on and off and at a lowered position. Similar data are
given for the airplane with modified wing leading
edges. Selected tuft photographs are given for
NACA RM A52B13
THE EFFECT OF BLUNTNESS ON THE DRAG OF
SPHERICAL-TIPPED TRUNCATED CONES OF
FINENESS RATIO 3 AT MACH NUMBERS 1.2 TO
7.4. Simon C. Sommer and James A. Stark. April
1952. 18p. photos., diagrs. (NACA RM A52B13)
(Declassified from Confidential, 10, 12/54)
The drag of spherically blunted conical models of
fineness ratio 3 was investigated in the Ames super-
sonic free-flight wind tunnel at Mach numbers from
1.2 to 7.4 in the Reynolds number range from
1.0 x 106 to 7.5 x 106. The models tested had blunt-
ness ratios of nose diameter to base diameter from
0 to 0.50. The use of small amounts of bluntness for
minimizing drag and the drag penalties associated
with large bluninesses are discussed.
NACA RM A52C20
DRAG OF CIRCULAR CYLINDERS FOR A WIDE
RANGE OF REYNOLDS NUMBERS AND MACH
NUMBERS. Forrest E. Gowen and Edward W.
Perkins. June 1952. 26p. diagrs., photos. (NACA
RM A52C20) (Declassified from Confidential,
Pressure distribution measurements on a two dimen-
sional circular cylinder have been made at high sub-
sonic and supersonic speeds. Drag coefficients ob-
tained from these measurements are presented with
results from other sources. It was found that a max-
imum drag coefficient of about 2.1 occurs near sonic
velocity. As the Mach number was increased to 2.9
the drag coefficient decreased to about 1.34. No e[-
fects of Reynolds number were found at supercritical
Mach numbers. Effects of fineness ratio on drag of
three-dimensional cylinders at supersonic speeds
were investigated and found to be small.
NACA RM A52C24
INVESTIGATION OF LIFT AND CENTER OF PRES-
SURE OF LOW-ASPECT-RATIO, CRUCIFORM,
TRIANGULAR, AND RECTANGULAR WINGS IN
COMBINATION WITH A SLENDER FUSELAGE AT
HIGH SUPERSONIC SPEEDS. Thomas N. Canning
and Billy Pat Denardo. June 1952. 28p. diagrs.,
photos. (NACA RM A52C24) (Declassified from
Confidential, 10 12. 54)
Low-aspect-ratio triangular and rectangular wings
in combination with a slender body have been tested
in the Mach number range between 1.3 and 6.2 and
Reynolds numbers based on body length from 2.8 to
16 million in the Ames supersonic free-flight wind
tunnel. The experimental lift-curve slope and cen-
ter of pressure positions are compared with theore-
tical predictions. Drag of the configurations is also
presented. The possible lift-drag ratios for the con-
figurations tested are estimated for Mach number
NACA RM A52DOla
EFFECTS OF PROPELLER-SPINNER JUNCTURE
ON THE PRESSURE-RECOVERY CHARACTERIS-
TICS OF AN NACA 1-SERIES D-TYPE COWL IN
COMBINATION WITH A FOUR-BLADE SINGLE-
ROTATION PROPELLER AT MACH NUMBERS UP
TO 0.83 AND AT AN ANGLE OF ATTACK OF 0.
Robert I. Sammonds and Ashley J. Molk. June 1952.
45p. diagrs., photos., Lab. (NACA RM A52DOla)
(Declassified from Confidential, 10, 12,54)
Measurement of ram-recovery ratio was made in the
duct of a cowling-spinner combination. Tests were
conducted at Mach numbers from 0.20 to 0.83 with
the propeller operating with an "ideal" propeller-
spinner juncture (propeller blade extended to spin-
ner surface) and a platform juncture (fixed airfoil-
shaped land that permitted blade angle change), and
with the propeller removed. Tests were run at pro-
peller blade angles of 600, 500. and 400, for
propeller advance-diameter ratios from 1.3 to 4.5
and inlet-velocity ratios from 0.26 to 1.33. All of
the tests were conducted at an angle of attack of 00
and a Reynolds number of 1.77 million, based on the
maximum diameter of the cowl.
NACA RM A52D11
THE TRANSONIC CHARACTERISTICS OF 38 CAM-
BERED RECTANGULAR WINGS OF VARYING AS-
PECT RATIO AND THICKNESS AS DETERMINED BY
THE TRANSONIC-BUMP TECHNIQUE. Warren H.
Nelson and Walter J. Krumm. July 1952. 173p.
diagrs., photos. (NACA RM A52D11) (Declassified
from Confidential, 10/12/,'54)
An investigation was made in the Ames 16-foot luhigh-
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 I, 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.
NACA RM A52D17
EFFECT OF TRAILING-EDGE THICKNESS ON LIFT
AT SUPERSONIC VELOCITIES. Dean R. Chapman
and Robert H. Kester. July 1952. 24p. diagrs..
photos. (NACA RM A52D17) (Declassified from
Confidential, 1012, 54)
Lift forces on various rectangular-plan-form wings
were measured in the Mach number range between
1.5 and 3.1 at Reynolds numbers between 0.55 and
2.2 million. The wings differed in trailing-edge
thickness, profile shape, maximum thickness ratio,
and aspect ratio. Measurements were made on
wings with and without a boundary-layer trip and are
compared to theoretical calculations. Calculated
results using shock-expansion theory are presented
for Mach numbers up to 10. In general, thickening
the trailing edge resulted in an increase in lift-
curve slope. This increase varied between a few
percent and about 15 percent, depending primarily,
on the trailing-edge thickness. Calculations indicate
that somewhat greater increases are possible at
high supersonic Mach numbers.
-ACA RM A52D24
t. IEMENTARY NOTE ON MODIFIED-IMPACT-
)I'm CALCUiLATIONS FOR BODIES OF REVO-
TN HAVING MlINIMUM DRAG AT HYPERSONIC
I:IS.Y Mey' M. llResniktoff. July 1952. 13p.
(NACA R'A A52D24) (Declassified from
Cc -i-l i' iIa I/' 4,
tnaa inp&-t `teory as modified in NACA
". 5 Ii, .l: 'o include an estimate of centrifu-
is us d to determine bodies for hypersonic
i having :imini pressure foredrag under vari-
i binatio"s f the conditions of given body
.i, 'i;pet, & olxne, and surface area.
4 E. H IRIMENf rA.. IrVESTIGATION OF THE
'i*'FRESSUR F "FI2ACTERISTICS OF NON-
1 BOLQE O F" \EVOLUTION AT MACH
4 I f:aS yRO.O.; 2'73 TO 4.98. John 0. Reller, Jr.
.I r rank '1. H1 akr. September 1952. 47p.
photo ;/CA TM A'-;r:2? (Declassified
fl-m alorfitneli al '1C/154)
'I ; rss re L C ar 'toistics of related nonlifting
f re voli re inve stigated at free-stream
I! nu.iiE rA ho ..3 tL 4.98 and Reynolds num-
i' *(r 0.6 x l '.8 1. The basic body
.t~i i'E i a It 'tive with a cylind-
'rt'od n'.iauon rf base pressure
'i t'.it -'!ir Mach number and
'*';. III!' !C miii ed for laminar-,
> it' 0a -, i: t -bound.:ry-layer flow.
*" 'It" ls i ['*'.\ i SS ratio, -" Il.ul'lre
." >.t'! 1 'U a^LV4 o 'D INGE-MOMENT
ii.'EPI S'h 0 SPEED OF LARGE-
i" I-. FT I A ED, FLAP-TYPE CON-
T '. AIL. Cl A! ING OF ASPECT
J: '1 : August 1952. 53p.
olC. NALA RM A52F13).
Iat, 10/ : li
C -* !i .I ige-nmoinent characteristics
j i,, .a xinC of aspect ratio 2
S i. pt i-otiols vithli a swept-
i. : rios sizes and shapes of
1,;. o'1f changes n the follow-
S.,r V g I thie horn balance,
*u *' oI; l 1 'Jdec thickness,
i>.'. Iint the corn balance, and
*1. C, ,' Itrols which had a
'a I 1 ) at sp.e-d was cornoared
t'. a' ;if1; i t*r s on triangular
>.' t '.. .I iagular-wing con-
I'i.GAri I; ACA SUBMERGED INLET
A'r '-. ijIB- iM 1.17 to 1.99. Warren
r'. a Frazer. September 1952.
29p. : (ACA RM A52F17)
(Decl, -.ii-t1' .> 'itial, 10/12/54)
RESEARCH ABSTRACTS NO. 73
An investigation was conducted with a small-scale
NACA submerged inlet at supersonic Mach numbers
from 1.17 to 1.99. The measured performance of a
submerged inlet at low supersonic Mach numbers
was compared to the calculated performance of a
normal-shock scoop inlet on the basis of net thrust
NACA RM A52I17
LATERAL AND DIRECTIONAL DYNAMIC-
RESPONSE CHARACTERISTICS OF A 350 SWEPT-
WING AIRPLANE AS DETERMINED FROM FLIGHT
MEASUREMENTS. William C. Triplett and
Stuart C. Brown. December 1952. 62p. diagrs.,
photo., 3 tabs. (NACA RM A52117) (Declassified)
from Confidential, 10/12/54)
Lateral and directional dynamic-response charac-
teristics, including frequency responses, transfer
functions, and stability derivatives of a 350 swept-
wing fighter airplane are obtained from flight
measurements of transient responses to rudder and
to aileron control deflections. Flight records were
obtained at two altitudes between Mach numbers of
0.50 and 1.04. Effects of aeroelasticity are dis-
cussed, and test results are compared to predictions
based on wind-tunnel and theoretical studies.
NACA RM A53G31
A CORRELATION BY MEANS OF TRANSONIC
SIMILARITY RULES OF THE EXPERIMENTALLY
DETERMINED CHARACTERISTICS OF 18 CAM-
BERED WINGS OF RECTANGULAR PLAN FORM.
John B. McDevitt. September 1953. 57p. diagrs.
(NACA RM A. 5' i. (Declassified from Confidential,
The effects of one type of camber on the aerodynamic
characteristics of rectangular wing at high subsonic
and transonic speeds have been studied by applying
the transonic similarity rules to the correlation of
experimental data for a series of 18 cambered wings
having NACA 63A2XX and 63A4XX sections, aspect
ratios from 1 to 4, and thicknesses from 4 to 8 per-
cent. The data were obtained by use of a transonic
bump over a Mach number range of 0.6 to 1.1.
NACA RM E51F15
INVESTIGATION OF DYNAMIC. CHARACTERISTICS
OF A TURBINE-PROPELLER ENGINE. Frank L.
Oppenheimer and James R. Jacques. September
1951. 22p. diagrs., tab. (NACA RM E51F15)
(Declassified from Confidential. 10/1 ". ij
Time constants that characterize engine speed re-
sponse of a turbine-propeller engine over the
cruising speed range ior various values of constant
fuel flow and constant blade angle were obtained both
from sleady-state characteristics and from transient
operating. Magnitude of speed response to change?
in fuel :flw and blade angle was investigated and is
presented in the form of gain factors. Results indi-
cate that at any given value of speed in the engine
cruising speed range, time constants obtained both
from steady-state characteristics and from transient
operation agree satisfactorily for any given constant
fuel flow, whereas time constants obtained from
RESEARCH ABSTRACTS NO. 73
transient operation exceed time constants obtained
l,ronn steady-stage characteristics by approximately
14 percent for any given blade angle.
NACA RM E51F19
PRELIMINARY INVESTIGATION OF THE CONTROL
OF A GAS-TURBINE ENGINE FOR A HELICOPTER.
Richard P. Krebs. September 1951. 13p. diagrs.
(NACA RM E51F19) (Declassified from Confidential,
10 12. 541
An analog investigation of the power plant for a gas-
turbine p-5aered helicopter indicates that currently
proposed turbine-propeller engine controls are
satisfactory for helicopter application. Power
increases from one-half to full rated at altitudes
from sea level to 15,000 feet could be made in less
than 4 seconds with either the rotor or propellers
aosurbing the engine power.
NACA RM E51J25
EXPERIMENTAL INVESTIGATION OF THE VIBRA-
TION CHARACTERISTICS OF FOUR DESIGNS OF
TURBINE BLADES AND OF THE EFFECT PRO-
DUCED BY VARYING THE AXIAL SPACING
BETWEEN NOZZLE BLADES AND TURBINE
BLADES. W. C. Morgan and C. R. Morse.
FEbruary 1952. 28p. diagrs., photos., tab. (NACA
FNI E5 1J25) (Declassified from Confidential,
10 1.? 54)
An investigation was made to determine the effects
of varying the spacing between the nozzle blades and
the turbine blades of a turbo-jet engine on turbine-
blaile vibration for four turbine-blade designs of
Jdii rent degrees of stiffness. In general, there
wais a tendency toward increase in occurrence of
vi-rjtin with decrease in spacing. The effect was
most evidiiin in the case of the turbine blades that
icid ureiter stiffness.
NAL A RM C51K21
A METHOD FOR ESTIMATING SPEED RESPONSE
OF GAS- TURBINE ENGINES. Harold Gold and
S'i.m)nn Hosenzweig. January 1952. 26p. diagrs.
(NACA RM E51K21) (Declassified from Con..rriderti.l.
10 12 541
A 'rie: nit itod is presented for estimating the speed
?a ep..nE 1.i turbojet and turbine-propeller engines
v, step change in fuel flow. The method approxi-
Imais th P dynamic equilibrium in the gas-turbine
e-rnginrie wiin a first-order linear differential equation,
the time constant of which varies inversely with the
eqiriiri un-i speed. The deviation of the calculated
values iron the mean experimental values is only
sli]hti,' greater than the spread of experimental
NACA RM E52B14
RELATIONS BETWEEN FUEL PROPERTIES AND
COMBUSTION CARBON DEPOSITION. Edmund R.
J, nash, Jcrrinld D. Wear and Robert R. Hibbard.
April 1i52. 67p. diagrs., 3 tabs (NACA
RMII E52B14) (Derlassified from Conideritial
10 I1 54)
Methods for pre'i- .io carbon-fornI.g propensity (A
turbojet-engine fuels from results t simple labora-
tory tests are discussed witn a viaw toward their
application to the control of jet fuel quality. The
methods involve the use of aromatic content, hydro-
gen carbon ratio, distillation temperatures, .. .m.
aniline point, and several empirical laboratory car-
bon deposition tests. Most accurate prediction of
carbon deposition was obtained with (1) the NACA K
factor (function of hydrogen -c arbon ratio and volu-
metric average boiling temperature), and (2) the
smoking tendency of the fue alt'iouh this latter
method requires addition t d.a t establish re-
producibility of the method among laboratories,
NACA RM E52DO10
INVESTIGATION OF ENGINE PERFORMANCE AND)
HIGH-TEMPERATURE PROPERTIES OF
PRECISION-CAST TURBINE BLADES OF H1I1-
CARBON SATELLITE 21 AND CONTRCLI.ED-GRAlI-
SIZE STELLITE 21. Charles Y ker, Floyr B.i
Garrett and Paul F. Sikora. June 1,52i. 31p.
diagrs., photos., 6 tabs. '.-ACA RM E51D10)
(Declassified from Confidential, 1 )'12, )
The effect of controlled :r.a 1 sio and increasedd ar-
bon content on engine pfera fa,.nce- and ui h e.npl In
ture properties of precis.oi-caI elIte 21 ir)om
blades was ir.vtz inii Blades ;e."e cas tto i..O-
in controlled grain sizes, cashed k. n 1 dia1, 4.11
coarse; others were cast o iig-' cI' n Stel. I'
The blades were operated i;a tu ri je engiie.
oratory stress-rupture testc spcinens cl rom
blades were conducted at t1i3 stre i n. I teuipe'alirr
encountered during rated. s eed e ini operation.
The results of the e-gi a lbor t ry teats intL
cated that (1) time 1 a it il es lad
lower life and crv; i .ii
and coarse-grain olaide a I I ir. I iti
blades showed less after wui 'u
blades; and (2) hghe r-ur *; 3 ;
bhlade life than t .aduz: Si'
NACA RM E52L1!
BEHAVIOR OF FO( CF! *s- mt E BFI)Ii
IN STEADy-STATF ICPEE 'S 3 ? 1 '1 f
JET ENGINE T S, rS-4im FU 4'.D!)
METALLOGRAliIC EVA iI. I S. t,
C. A. C2.. i .h ar i. W. vi. 'i r
29p. diagrs., ph'ios 3 t&i: (A. EM i_.
mIV'i 1 .I*-*im,-. from Confrden.til, 'Ci?,
An ; in .tii, *.. 1as conduct- ro i0 (m 'ne -
:avior of pr due f:' 0- 1 "I'.
blades in a fui Io turbt C f .. i0 ir-
ular, the scatter in perrorina> -
turbine blades were operated a i.A a ." *
sible at a tempat iaurme oft bl' .. .'
stress of lr. pou-ds peri t
.1-12ives 'M the turbine I' r J ) i
539 hours, arange of 38 hour. r' .p!
'*'run .ii'.i- of sp cimens a ;f'o' l~r. i:ft~ -***
varied considerably, as im..X a
20,000 pounds per square ia .
the variability or scatter ( te'> .
greater than that of blade p;. e
is probably caused by varian :e .
the forged blades rather tha
by engine operation or inst. .
Metallographic examinations a ,
mine possible causes of the scatter and although
numerous differences in microstructures of blades
were found, no consistent trendencies were observed
and the findings did not permit an explanation of the
scatter of blade performance. The results of the
metallographic examinations and of the physical
tests indirectly indicated variables in the fabricating
method caused the scatter in properties.
NACA RM E52L30
ADHESIVE AND PROTECTIVE CHARACTERISTICS
OF CERAMIC COATING A-417 AND ITS EFFECT
ON ENGINE LIFE OF FORGED REFRACTALOY-26
(AMS 5760) AND CAST SATELLITE 21 (AMS 5385)
TURBINE BLADES. Floyd B. Garrett and Charles
A. Gyorgak. February 1953. 21p. photos., diagr.,
4 tabs. (NACA RM E52L30) (Declassified from
The adhesive and protective characteristics of
National Bureau of Standards Coating A-417 were
investigated, as well as the effect of the coating on
the life of forged Refractaloy 26 and cast Stellite 21
turbine blades. Coated and uncoated blades were
run in a full-scale J33-9 engine and were subjected
to simulated service operations consisting of con-
secutive 20-minute cycles (15 mmin at rated speed and
approximately 5 min at idle). The ceramic coating
adhered well to Refractaloy 26 and Stellite 21 turbine
blades operated at 15000 F. The coating also pre-
vented corrosion of the Refractaloy 26, a corrosion-
sensitive nickel-base alloy, and of the Stellite 21,
a relatively corrosion-resistant cobalt-base alloy.
Although the coating prevented corrosion of both
alloys, it had no apparent effect on blade life.
NACA RM E53A19
COMPARISON OF THEORETICALLY AND EXPERI-
MENTALLY DETERMINED EFFECTS OF OXIDE
COATINGS SUPPLIED BY FUEL ADDITIVES ON
UNCOOLED TURBINE-BLADE TEMPERATURE
DURING TRANSIENT TURBOJET-ENGINE OPERA-
TION. Louis J. Schafer, Jr., Francis S. Stepka
and W. Byron Brown. March 1953. 45p. photos.,
diagrs., tab. (NACA RM E53A19) (Declassified
from Confidential, 10/12/54)
An analysis was made to permit the calculation of the
effectiveness of oxide coatings in retarding the tran-
sient heat flow into turbine blades when the combus-
tion gas temperature of a turbojet engine is suddenly
changed. The analysis is checked with experimental
data obtained from a turbojet engine whose blades
were coated with two different coating materials
(silicon dioxide and boric oxide) by adding silicone
oil and tributyl borate to the engine fuel. The very
thin coatings (approximately 0.001 in.) that formed
on the blades produced a negligible effect on the
turbine-blade transient temperature response.
From the analysis of this report it was possible to
predict the turbine rotor-blade temperature response
with a maximum error of 400 F.
NACA RM E53E22
BURNING RATES OF SINGLE FUEL DROPS AND
THEIR APPLICATION TO TURBOJET COMBUS-
TION PROCESS. Charles C. Graves. July 1953.
35p. diagrs., photos., tab. (NACA RM E53E22)
(Declassified from Confidential, 10/12/54)
RESEARCH ABSTRACTS NO. 73
Burning rates were determined for single isooctane
drops suspended in various quiescent oxygen-
nitrogen atmospheres at room temperature and
pressure. The burning rates were compared with
those predicted by a previously developed theory
based on a heat- and mass-transfer mechanism and
with values predicted by a modification to this
theory. The drop-burning-rate data were applied to
equations for a burning fuel spray in order to calcu-
late the predicted change in burning rate of a fuel
spray with variation in oxygen concentration. The
results so obtained were compared with the change
in combustion efficiency of a single turbojet combus-
tor with inlet oxygen concentration, as determined
in a previous investigation. The drop burning rates
were proportional to drop diameter and increased
approximately 34 percent when oxygen concentration
of the surrounding oxygen-nitrogen atmosphere was
raised from 17.0 to 34.9 percent by volume. The
experimentally determined burning rates agreed
well with those predicted by the modified heat- and
mass-transfer theory. The predicted change in
combustion efficiency with inlet oxygen concentra-
tion was appreciably smaller than that observed in
the combustor tests.
NACA RM L51E24a
BUFFETING-LOAD MEASUREMENTS ON A JET-
POWERED BOMBER AIRPLANE WITH REFLEXED
FLAPS. John A. See and William S. Aiken, Jr.
August 1951. 28p. diagrs., 3 tabs. (NACA
RM L51E24a) (Declassified from Confidential,
Buffet boundaries, buffeting-load increments for the
stabilizers and elevators, and buffeting bending-
moment increments for the stabilizers and wings as
measured in gradual maneuvers for a jet-powered
bomber airplane equipped with reflexed flaps and
ailerons and tail tip incidence changes are presented
and compared with similar results for the original
airplane configuration. The Mach numbers of the
tests ranged from 0.35 to 0.81 at pressure altitudes
close to 30,000 feet. The predominant buffeting
frequencies were close to the natural frequencies of
the structural components. The magnitudes and
trends of buffeting-load coefficients witn Mach
number for the reflexed-flap configuration were
similar to those for the original configuration.
NACA RM L51G13
AN INVESTIGATION OF PROPELLER VIBRATIONS
EXCITED BY WING WAKES. W. H. Gray and
William Solomon. January 1952. 31p. diagrs.,
photo., tab. (NACA RM L51G13) (Declassified
from Confidential, 10/12/54)
This paper shows the effect of airspeed and wing
drag on the magnitude of vibratory stresses experi-
enced by a propeller operating in a wing wake. It
establishes the linearity of the relation between the
stresses and the velocity at a fixed value of wing
drag coefficient and between stress and drag coeffi-
cient at a fixed velocity. It also shows that an
augmented stress may be experienced by a propeller
operating too close behind the wing trailing edge.
RESEARCH ABSTRACTS NO. 73
N.CA RM L51112
PRESSURE PULSATIONS ON RIGID AIRFOILS AT
TRANSONIC SPEEDS. Milton D. Humphreys.
December 1951. 21p. diagrs., photos., tab.
(NACA RM L51112) tDeclassified from Confidential,
10, 12. 54)
The effect of changes in Mach number, thickness
ratio, and angle of attack on the amplitude of the in-
stantaneous pressure pulsations acting on airfoils
ranging in thicknesses from 4 to 12 percent chord
has been obtained at transonic speeds, and the cor-
responding flows past the airfoils were recorded by
high-speed schlieren motion pictures. The results
indicate that reduction in airfoil thickness was ac-
companied by marked reductions in the severity of
the aerodynamic pressure pulsations which contrib-
ute to airplane buffeting.
NACA RM L51J24
THE UNSYMMETRICAL LOAD AND BENDING
MOMENT ON THE HORIZONTAL TAIL OF A JET-
POWERED BOMBER MEASURED IN SIDESLIPPPING
FLIGHT. T. V. Cooney. January 1952. 19p.
diagrs., tab. (NACA RM L51J24) (Declassified
from Confidential, 10, 12/54)
Results are presented of an analysis of unsymmet-
rical tail-load and bending-moment measurements
made during sideslipping flight tests of a jet-powered
bomber airplane having a horizontal tail with 120 of
geometric dihedral. The unsymmetrical loads and
moments due to dihedral and to induction effects from
the vertical tail are deduced from the measurements.
A comparison of the results with estimates based on
available theory was found to be favorable, suggest-
ing that existing methods might be used to modify
present arbitrary design requirements relating to
unsymmetrical flight conditions.
NACA RM L51K08
LOW-SPEED STATIC LONGITUDINAL STABILITY
AND CONTROL CHARACTERISTICS OF 600
TRIANGULAR-WING AND MODIFIED 600
TRIANGULAR-WING MODELS HAVING HALF-
DELTA AND HALF-DIAMOND TIP CONTROLS.
Jacob H. Lichlenstein and Byron M. Jaquet.
February 1952. 36p. diagrs., photos., tab. (NACA
RM L51K08) (Declasstiied from Confidential,
The results are presented of an investigation made
in the Langley stability tunnel to determine the low-
speed static longitudinal stability and control effec-
tiveness characteristics of a basic and a modified
600 triangular-wing model having 10-percent-wing-
area half-delta and half-diamond, and 5-percent-
wing-area half-diamond tip controls. Comparisons
are made between the half-delta and half-diamond
controls and between the basic and modified
NACA RM L51K19
WIND-TUNNEL INVESTIGATION OF A SHIELDED
TOTAL-PRESSURE TUBE AT TRANSONIC SPEEDS.
William Gracey, Albin 0. Pearson and Walter R.
Russell. January 1952. 8p. diagrs. (NACA
RM L51KI9) (Declassified from Confidential,
The variation of total-pressure error with angle of
attack of a shielded total-pressure tube having a
curved venturi entry has been determined through an
angle-of-attack range of 00 to 600 at Mach numbers
ranging from 0.90 to 1.10. The results showed that
the tube measured total pressure correctly (to within
1 percent of the impact pressure) for angles of attack
up to about 570 at a Mach number of 0.90 and 560 at
a Mach number of 1.10.
NACA RM L51K23
THE INTERFERENCE EFFECTS OF A BODY ON
THE SPANWISE LOAD DISTRIBUTIONS OF TWO
450 SWEPTBACK WINGS OF ASPECT RATIO 8
FROM LOW-SPEED TESTS AT A REYNOLDS
NUMBER OF 4 x 106. Albert P. Martina. February
1952. 48p. diagrs., photo., 2 tabs. (NACA
RM L51K23) (Declassified from Confidential,
Tests of two wing-body combinations have been con-
ducted in the Langley 19-foot pressure tunnel at a
Reynolds number of 4 x 106 and a Mach number of
0.19 to determine the effects of the bodies on the
wing span load distributions. The wings had 450
sweepback of the quarter-chord line, aspect ratio
8.02, and taper ratio 0.45. One wing was untwisted
and incorporated NACA 631A012 airfoil sections in
the streamwise direction; the second wing was
twisted and cambered. The wings were mounted In
mid-high-wing positions on identical bodies of revo-
lution of 10:1 fineness ratio having maximum diam-
eters of 10 percent of the spans. The effects on the
incremental loading due to the body resulting from
wing incidence, upper-surface wing fences, and
flap deflection were also determined for the plane
wing. The body effects as calculated by using
several existing methods are compared with the
NACA RM L51K28
LONGITUDINAL FREQUENCY-RESPONSE CHAR-
ACTERISTICS OF THE DOUGLAS D-558-I AIR-
PLANE AS DETERMINED FROM EXPERIMENTAL
TRANSIENT-RESPONSE HISTORIES TO A MACH
NUMBER OF 0.90. Ellwyn E. Angle and Euclid C.
Holleman. February 1952. 28p. diagrs., tab.
(NACA RM L51K28) (Declassified from Confidential,
Transient responses from elevator pulses of the
Douglas D-558-I research airplane are analyzed by
the Fourier transform to give the longitudinal fre-
quency response of the airplane to a Mach number of
0.90 of altitudes between 30,000 and 37,000 feet.
NACA RM L51K28a
FLUTTER INVESTIGATION OF TWO THIN, LOW-
ASPECT-RATIO, SWEPT, SOLID, METAL WINGS
IN THE TRANSONIC RANGE BY USE OF A FREE-
FALLING BODY. W. T. Lauten, Jr. and Maurice
A. Sylvester. February 1952. 12p. diagrs., photo.,
2 tabs. (NACA RM L51K28a) (Declassified from
Two thin, interceptor-type (low-aspect ratio), 450
sweptback, untapered wings of solid metal construc-
tion have been tested for flutter up to a Mach number
of 1.23 by the free-falling-body technique. No flutter
was obtained. Flutter calculations yielded results
which showed that two-dimensional compressible
flow theory is not adequate for predicting flutter for
this type of low-aspect-ratio wing in the transonic
NACA RM L51L04
WIND-TUNNEL INVESTIGATION AT HIGH AND LOW
SUBSONIC MACH NUMBERS OF A THIN SWEPT-
BACK WING HAVING AN AIRFOIL SECTION DE-
SIGNED FOR HIGH MAXIMUM LIFT. Stanley F.
Racisz and Nicholas J. Paradise. February 1952.
46p. diagrs., photo., tab. (NACA RM L51L04)
(Declassified from Confidential, 10/12/54)
An investigation has been made of a semispan wing
with 450 sweepback, aspect ratio 4, and taper ratio
0.6, equipped with an airfoil section designed to
have high maximum lift at low Mach numbers. The
lift, drag, and pitching-moment characteristics
were determined at Reynolds numbers ranging from
2 x 106 to 9 x 106 for Mach numbers below 0.2 for
the wing with and without a split flap. The charac-
teristics of the plain wing were also determined for
several values of the Reynolds number at Mach num-
bers up to about 0.95. The results are compared
with results obtained from tests of a wing with the
same plan form but with the NACA 65A006 airfoil
NACA RM L51L11
AERODYNAMIC CHARACTERISTICS AT TRANSONIC
SPEEDS OF A TAPERED 450 SWEPTBACK WING
OF ASPECT RATIO 3 HAVING A FULL-SPAN FLAP
TYPE OF CONTROL WITH OVERHANG BALANCE.
TRANSONIC-BUMP METHOD. Vernard E.
Lockwood and John R. Hagerman. January 1952.
24p. diagrs. (NACA RM L51L11) (Declassified
from Confidential, 10/12/54)
Lift, pitching-moment, rolling-moment, and flap
hinge-moment coefficients were obtained by the
transonic bump method on a 45. sweptback wing of
aspect ratio 3, and taper ratio of 0.5, and an NACA
64A010 section, employing a 0.254-chord full-span
RESEARCH ABSTRACTS NO. 73
flap. The flap had a 50-percent flap-chord
elliptical-nose overhang. The investigation was
made at angles of attack of -40, 00, 40, and 80, flap
deflections from -280 to 60, and Mach numbers
from 0.6 to 1.15. The results are compared with
the same flap without overhang.
NACA RM L51L14
LOW-SPEED TESTS OF A FREE-TO-YAW MODEL
IN TWO WIND TUNNELS OF DIFFERENT TURBU-
LENCE. Jones F. Cahill and John D. Bird.
February 1952. 12p. diagrs., photos. (NACA
RM L51L14) (Declassified from Confidential,
Tests have been made at low speeds in the Langley
low-turbulence pressure tunnel which has a very low
turbulence level and the Langley stability tunnel
which has a turbulence level approximately ten times
as great in order to determine the extent of any re-
sulting oscillations of a model mounted with freedom
in yaw and in order to demonstrate the extent to
which directional fluctuations in an air stream can be
responsible for such oscillations. Tests covered a
range of dynamic pressures from 4 to 175 pounds per
square foot with associated ranges of Reynolds num-
bers and Mach numbers from 1.5 x 106 to 4.6 x 106
and 0.05 to 0.34, respectively.
NACA RM L51L19
PRELIMINARY INVESTIGATION OF THE EFFECTS
OF A PADDLE BALANCE ON THE CONTROL
CHARACTERISTICS AT TRANSONIC SPEEDS OF A
TAPERED 45.580 SWEPTBACK WING OF ASPECT
RATIO 3 HAVING A FULL-SPAN FLAP-TYPE CON-
TROL. William C. Moseley, Jr. February 1952.
24p. diagrs., photo. (NACA RM L51L19) (Declas-
sified from Confidential, 10/12/54)
A preliminary investigation at transonic speeds was
made on a wing having a quarter-chord line swept
back 45.580, an aspect ratio of 3, a taper ratio of
0.5, and an NACA 64A010 airfoil section. The
Reynolds number of the tests was approximately
1, 000, 000. The wing had a full-span flap tested
with and without a paddle balance. The paddle
balance was capable of balancing excessive control
hinge moments with only slight effects on the lift and
NACA-Langley 11-9-54 4M
Item Q~uantity Code Title and Author NACA
No. desired number (Only Needed When Code Number Unavailable) Action
POLICY OF NACA ON DISTRIBUTION OF THEIR PUBLICATIONS
NACA Reports. Technical Notesi, and Technical Memorandums are available for a period of 5 years,
after that, most of them. can be had only on a loan basis. All Wartime Reports are in this category.
All loan material should be returned promptly at the expiration of the loan period to the following address:
Langley Aeronautical Laboratory, Langley Field, Virginia ATTE;NTION: Mr. Walter H. Lee.
Brrtlsh publications currently listed on the Research Abstracts are available only on loan. However,
should a British paper be of particular interest and if you will so advise this office, your name will be
placed on our waiting list to receive a copy if and when retention copies can be furnished.
Please fill to the requested information below since the above part of this form will be returned with the
Date 19q Do Not Write in This Space
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
1512 H Street, N.W. Washington 25, D. C.
Division of Research Information
The Committee is pleased to forward the enclosed pub-
lications in accordance with your recent request.
~We regret that the remaining items are not enclosed for
the reasons) indicated.
A._ Out of print.
B. _Will supply when
C. Not an NACA document.
D. _Available on loan only.
E. _Photocopies available at
Library of Congress.
F. Cannot identify docu-
G. _Clas sified document.
Request through mili-
tary project officer.
H Withdrawn from cir-
1. -Not available for cir-
Documents on loan to be returned by
Cuty, Zone No., and State
City, Zone No.. and State
Ri~QEE OUR S
Do Not Write in This Space
Street address _____
City, Zone No.. and State
UNIVERSITY OF FLORIDA
3 1262 08153 096 5
xml version 1.0 encoding UTF-8
REPORT xmlns http:www.fcla.edudlsmddaitss xmlns:xsi http:www.w3.org2001XMLSchema-instance xsi:schemaLocation http:www.fcla.edudlsmddaitssdaitssReport.xsd
INGEST IEID EQT44NDJN_66R0JM INGEST_TIME 2012-03-02T22:20:13Z PACKAGE AA00009235_00036
AGREEMENT_INFO ACCOUNT UF PROJECT UFDC