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:
AA00009235:00037

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Full Text

'AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW, WASHINGTON 9s, D. C.,OTING CODE NUMBER ABOVE EACH TITLts
THE REPORT TITLE AND AUTHOR.


NACA Rept. 1163

A VISUALIZATION STUDY OF SECONDARY FLOWS
IN CASCADES. Howard Z. Herzig, Arthur G.
Hansen and George R. Costello. 1954. ii, 51p.
diagram photos. (NACA Rept. 1163. Formerly
TN 2947; RIM E52F19)

Flow-visualization techniques are employed to ascer-
tain the streamline patterns of the nonpotential, sec-
ondary flows in the boundary layers of cascades,
thereby providing a basis for more extended analyses
in turbomachines. The three-dimensional deflection
of the end-wall boundary layer results in the forma-
tion of a vortex well up in each cascade passage. The
size and tightness of the vortex generated depend
upon the main flow turning in the cascade passage.
Once formed, a vortex resists turning in subsequent
blade rows. This results in unfavorable angles of
attack and possible flow disturbances on the pressure
surfaces of subsequent blade rows when the vortices
impinge on these surfaces. Two major tip-
clearance effects are observed: the formation of a
tip-clearance vortex, and the scraping effect of a
blade with relative motion past the wall boundary
layer. The flow patterns Indls Inquiods~or im-
proving the blade-tip load I C~rcharaciericsof com-
pressors and of low- and kighydpeed turbidyl;. '


NACA Rept. 1164 1:sDi 315

CONVECTION OF A PATTERN OF VORTIC)TY
THROUGH SHOCK WAPE. R.S. Rlbnor. LO54.
ii, 17p. diagrs. (NACA Rept. 41'64. Formpflyf
TN 2864) .

An arbitrary weak spatial distribution of vorticity
can be represented in terms of plane sinusoidal
shear waves of all orientations and wave lengths
(Fourier integral). The analysis treats the passage!
of a single representative weak shear wave through
a plane shock and shows refraction and modification
of the shear wave with simultaneous generation of
an acoustically intense sound wave. Applications
to turbulence and to noise in supersonic wind tunnels
are indicated.


NACA Rept. 1171

EFFECT OF HORIZONTAL-TAIL SPAN AND
VERTICAL LOCATION ON THE: AERODYNAMIC
CHARACTERISTICS OF AN UNSWEPT TAIL
ASSEMBLY IN SIDESLIP. Donald R. Riley. 1954.
it, 20p. diagrs., photos., tab. (NACA Rept. 1171.
Formerly TNJ 2907)

Wind-tunnl results on the effect of hoirzontal-tail
span and vertical location of the horizontal tail
relative to the vertical tail on the aerodynamic


National Advisory Committee For Aeronautics


Research Abstracts rb
NO;VEMBER 30, 1954


NO. 74


CURRENT NACA REPORTS

NACA Rept. 1133

MlECHANISM OF START AND DEVELOPMENT OF
AIRCRAFT CRASH FIRES. I. Irving Pinkel, G.
Merritt Preston and Gerald J. Pesman. 1953. iii,
52p. diagrs., photos., 2 tabs. (NACA Rept. 1133.
Formerly RM E52F06)

Full-scale aircraft crashes were made to investigate
the mechanism of the start and development of air.
craft crash fires. The results are discussed herein.
This investigation revealed the characteristics of the
ignition sources, the manner in which the combusti-
bles spread, the mechanism, of the union of the com-
bustibles and ignition sources, and the pertinent
factors governing the development of a crash fire as
observed in this program,

NACA Rept. 1134

PHOTOGAPHIC INVJTESTIGATION OF COMBUSTION
IN AP TWO-DIMENSIONAL TRANSPARENT ROCKET
ENGIN. Donald R. Bellman, Jack C. Humphrey
and Theodore Male. 1953. ii, 12p. diagrs., photos.,
tab. (NACA Rept. 1134. Formerly RM E8F01)

Motion pictures at camera speeds up to 3000 frames
per second were taken of the combustion of liquid
oxygen and gasoline in a 100-pound thrust rocket
engine. The effect of seven methods of propellant
injection on the uniformity of combustion was in-
vestigated. The flame front was generally found to
extend to the injector faces and all the injection
systems showed considerable nonuniformity of
combustion. Pressure vibration records indicated
combustion vibrations that corresponded to
resonant-chamber frequencies,


NACA Rept. 1151

THEI EFFECTS ON DYNAMIC LAERAL STABILITY
AND) CONTROL OF IARGjE ARTIFICIAL VARIA-
TIONS IN THE ROTARY STABIIY DERIVATIVES.
Robert O. Schade and James L. Hfassell, Jr. 1953.
ii, 24p. diagrs., photo., 2 tabs. (NACA Rept. 1151.
Forme rly TN 2781)

The results of an experimental and theoretical inves-
tigation of the effects of large artificial variations
of four rotary stability derivatives on the dynamic
lateral stability and control of a 45o sweptback-wing
airplane model are presented. The experimental
results are presented mainly in the form of flight
ratings for stability, control, and general flight be-
havior. Calculations at period and dating and of
the response to rolling and yawing disturbances are
also presented.









characteristics of an unswept vertical-tail assembly
in sideslip are presented. By applying the well-
known discrete-horseshoe-vortex method used for
wings to the problem of intersecting surfaces,
theoretical span loadings were obtained for each of
the configurations tested. Calculated values ob-
tained from the span loadings are compared with
experimental results.





NACA Rept. 1172

A STUDY OF THE APPLICATION OF POWER-
SPECTRAL METHODS OF GENERALIZED HARMON-
IC ANALYSIS TO GUST LOADS ON AIRPLANES.
Harry Press and Bernard Mazelsky. 1954. ii, 17p.
diagrs., 2 tabs. (NACA Rept. 1172. Formerly
TN 2853)

The applicability of some results from the theory of
generalized harmonic analysis to the analysis of gust
loads on airplanes in continuous rough air is exam-
ined. The input and output relations in terms of
power spectrums are used to relate the standard de-
viation (root mean square) of loads in continuous
rough air to the gust response characteristics of the
airplane and the spectral characteristics of atmos-
pheric turbulence. For the case of a normally dis-
tributed output, which appears from experimental
evidence to apply to a homogeneous turbulence condi-
tion, the probability distribution of loads is shown to
be a simple function of the output power spectrum*
In order to illustrate the application of power-
spectral analysis to gust loads and to obtain an n-
sight into the relation between loads in continuous
rough air and the gust response characteristics of the
airplane, two selected series of calculations are pre-
sented. The results of these applications are com-
pared with those obtained from calculations for single
gusts and their implications are discussed-




NACA TM 1366

HEAT TRANSFER BY FREE CONVECTION FROM
HORIZONTAL CYLINDERS IN DIATOMIC GASES.
(Warmetibergang bei breier Stromung am wagrechten
Zylinder in zweiatomigen Gasen). R. Hermann.
November 1954. 73p. diagrs., photos., 6 tabs.
(NACA TM 1366. Trans. from VDI Forschungsheft,
No.379, 1936)

Heat transfer by free convection from horizontal cyl-
inders in diatomic gases is investigated theoretically
and experimentally. The heat transfer is given by
the theoretical equation Nu = 0.37(Gr)1/4 for
104 < Gr < 3 x 108. Experimental determinations of
velocity, temperature, and heat transfer are in good
agreement with the theory. It is found that the total
heat transfer and the boundary-layer development at
the upper stagnation point and at the upper edge,
respectively, of a horizontal cylinder are equivalent
to the respective quantities for a vertical plate which
transfers heat on both sides and is 1.31 times the
cylinder diameter. A review and discussion of pre-
vious investigations of free-convection heat transfer
from horizontal cylinders is included.


NACA
RESEARCH ABSTRACTS NO.74


NACA TM 1373

ON FORCE-DEFLECTION DIAGRAMS OF AIRf-
PLANE SHOCK ABSORBER STRUTS. FIRST,
SECOND, AND THIRD PARTIAL REPORTS. (Zur
Kenntnis der Kraftwegdiagramme von
Flugzeugfederbeinen). K. Schlaefke. November
1954. 48p. diagrs., 4 tabs. (NACA TM 1373.
Trans. from Technische Berichte, v. 11, nos. 2, 4,
& 5, 1944)

This paper, which is presented in three parts, is an
analytical study of the behavior of landing-gear
shock struts, with various types of assumptions for
the shock-strut characteristics. The effects of tire
springing are neglected. The first part compares
the behavior of struts with linear and quadratic
damping. The second part considers struts with
nonlinear spring characteristics and linear or
quadratic damping. The third part treats the oleo-
pneumatic strut with air-compression springing
without damping and with damping proportional to
velocity. It is indicated how the damping factor can
be determined by experiment.






NACA TM 1378

DETERMINATION OF THE ELASTIC CONSTANTS
OF AIRPLANE TIRES. (Ermittlung der elastischen
Konstanten von Flugzengreifen). Boeckh. November
1954. 30p. diagrs. (NACA TM 1378. Trans. from
Focke-Wulf Flugzeugbau G.m.b.H, Bremen,
V.13.3703)

Measurements were made of the distortion of four
German aircraft tires, from about 22 to 28 inches
in diameter, at several vertical loadings. For each
vertical loading measurements were made of the
tire distortion under several lateral, tangential and
torsional loadfings.







NACA TM 1381

ON THE DETERMINATION OF CERTAIN BASIC
TYPES OF SUPERSONIC FLOW FIELDS. (Sulla
determinazione di alcuni tipi di campi di corrente
ipersonora). Carlo Ferrari. November 1954. 17p.
diagrs. (NACA TM 1381. Trans. from Rendiconti
della R. Accademie Nazionale dei Lincei, Series 8,
v.7, no. 6, Dec. 1949)

A discussion is given of the application of Fourier
series techniques to the problems of linearized super-
sonic flow. The formulation presented is an exten-
sion of the doublet type of "fundamental solution' to
higher order types of singularity. The equations
developed have application to wing theory but are pri-
marily of importance in wing-body interaction pro-
blems. A specific example of a wing-body interfer-
ence problem is discussed in light of the presented
methods.






NACA
RESEARCH ABSTRACTS NO. 74

NACA TN 3217

THE INFLUENCE OF WHEEL SPIN-UP ON
LANDING-GEAR IMPACT. W. Fliigge and C. W.
Coale, Stanford University. October 1954. ii, 107p.
diagrs., tabs. (NACA TN 3217. Continuation of
TN 2743)

This report deals with the influence of wheel drag on
landing-gear performance. The differential
equations are developed and solved by numerical
integration and by an analytical method. Emphasis
is placed on dropping influences of minor impor-
tance to simplify the computations. Consideration
is also given to the influence of offset wheels, in-
elined shock struts, and friction due to ovalization
of the shock-strut cylinder under bending loads.


NACA TN 3219

VISCOSITY CORRECTIONS TO CONE PROBES IN
RAREFIED SUPERSONIC FLOW AT A NOMINAL
MACH NUMBER OF 4. L. Talbot, University of
California. November 1954. 39p. diagrs.,photo.,
4 tabs. (NACA TN 3219)

Viscosity correction data were obtained for a set of
geometrically similar cone probes in supersonic
rarefied air flow at a nominal Mach number of 4.
Additional experiments were made to determine
cone surface pressure distributions and the effects
of variable orifice size on the measurement of cone
surface pressure. Pressure distributions near the
vertex were compared with the tangent-cone and
linearized theories.



NACA TN 3242

PRELIMINARY RESULTS FROM FLOW-FIELD
MEASUREMENTS AROUND SINGLE AND TANDEM
ROTORS IN THE LANGLEY FULL-SCALE TUNNEL.
Harry H. Heyson. November 1954. 19p. diagrs.,
photos. (NACA TN 3242)

Measurements of the flow angles and velocities near
lifting helicopter rotors, both single and tandem, at
a tip-speed ratio of 0.15 are presented and compared
with those of theory. The comparison indicates that
the theory is sufficiently accurate for use in prelim -
inary design calculations. The flow behind a rotor
is shown to be very much like the flow behind a wing.


NACA TN 3252

DESCRIPTION AND PRELIMINARY FLIGHT INVES.
TIGATION OF AN INSTRUMENT FOR DETECTING
SUBNORMAL ACCELERATION DURING TAKE-OFF.
Garland J. Morris and Lindsay J. Lina. November
1954. 19p. diagrs, photos. (NACA TN 3252)

An instruent actuated by longitudinal acceleration
and impact pressure has been proposed which would
give a quick and easily recognizable quatitative in-
dication of loss in airplane acceleration during take-
off. A preliminary devaluation from flight tests of a
simplified prototype instruent mounted in a jet
trainer has been made. The instrument was found
to be satisfactory and the response of the indicator
to simulated partial power loss was rapid.


NACA TN 3260

SMOKE STUDY OF NOZZLE SECONDARY FLOWS
IN A LOW-SPEED TURBINE. Mnilton G. Kofskey
and Hubert W. Allen. November 1954. 24p. diagrs.,
photos. (NACA TN 3260)

Still and motion pictures wvere made of boundary-
layer and wake secondary-flow phenomena visualized
by smoke. Two annular cascades of turbine nozzles
were used, both designed for constant discharge
angle but differing in blade shape and suction-
surface velocity distribution. Flows were similar
to those obtained with pressure and angle measure-
ments at near-sonic airspeeds. Boundary-layer
cross-channel and trailing-edge radial flows caused
vortices and an accumulation of low-momentumn air
at the hub, which may affect flow in following blade
rows. Motion of a downstream rotor blade row
produced pulsations in trailing-edge radial flow.
The motion-picture supplement may be obtained on
loan from NACA Headquarters, Washington, D. C.



NACA TN 3268

SHEARING-STRESS MEASUREMENTS BY USE OF A
HEATED ELEMENT. H. W. Liepmann and G. T.
Skinner, California Institute of Technology.
November 1954. 27p. diagrs. (NACA TN 3268)

The present report discusses the use of an instru-
ment to determine the local skin-friction coefficient
by measurement of heat transfer from small ele-
ments embedded in the surface. The range of
application of such instruments is discussed and
experimental data are presented to show that a
simple instrument consisting of an ordinary hot-
wire cemented into a groove in the surface can be
used to obtain laminar and turbulent skin-friction
coefficients with a single calibration.



NACA TN 3283

AERODYNAMIC FORCES, MOMENTS, AND STA-
BILITY DERIVATIVES FOR SLENDER BODIES OF
GENERAL CROSS SECTION. Alvin H. Sacks.
Novemberl954.(i), 74p. diagrs., 2 tabs. (NACA
TN 3283)

Formulas are developed for forces and moments in
terms of the body shape and motions, and for sta-
bility derivatives in terms of the mapping functions
of the cross sections. Relationships are found
among the various derivatives, and calculations are
made for several configurations. The influence of
the squared terms in the pressure relation is
demonstrated, and the use of the apparent mass
concept is discussed in detail.




NACA TN 3286

GENERALIZED INDICIAL FORCES ON DEFORMING
RECTANGULAR WINGS IN SUPERSONIC FLIGHT.
Harvard Lomax, Franklyn B. Fuller and Loma Sluder.
November 1954.74p. diagrs., tab. (NACA TN 3286)






NACA
RESEARCH ABSTRACTS NO. 74

NACA TN 3308

AN EXPLORATORY INVESTIGATION OF SOME
TYPES OF AEROELASTIC INSTABILITY OF OPEN
AND CLOSED BODIES OF REVOLUTION MOUNTED
ON SLENDER STRUTS. S. A. Clevenson,
E. Widmayer, Jr. and Franklin W. Diederich.
November 1954. 44p. diagrs., photos., 3 tabs.
(NACA T~N 3308. Formerly RM L53E07)

The aeroelastic instability of rigid open and closed
bodies of revolution mounted on thin, flexible struts
has been investigated experimentally at low speeds.
Three types of instability were observed coupled
flutter, divergence, and an uncoupled oscillatory in-
stability which consists in continuous or intermittent
small-amplitude yawing oscillations. An attempt has
been made to calculate the airspeeds and, in the case
of the oscillatory phenomena, the frequencies at which
tese t pest of instability ocur byusin a sled-body



NACA TN 3310

INVESTIGATION OF STATIC STRENGTH AND
CREEP BEHAVIOR OF AN ALUMINUM-ALLOY
MULT1WEB BOX BEAM AT ELEVATED TEMPER-
ATURES. Eldon E. Mathauser. November 1954.
21p. diagrs., photos., 4 tabs. (NACA TN 3310)

Results of an investigation to determine static
strength and creep behavior at elevated temperatures
of 24S-T3 aluminum-alloy multiweb box beams are
presented. Methods that were used to predict failure
stresses in the static-strength tests were in good
agreement with the experimental results. Creep
deflections and creep lifetimes are presented for
beams subjected to constant load and various heating
conditions. Lifetime is satisfactorily predicted from
material stress-rupture data when tensile failure
occurs at both constant and varying temperatures.


NACA TN 3313

SOME MEASUREMENTS OF ATMOSPHERIC
TURBULENCE OBTAINED FROM FLOW-DIRECTION
VANES MOUNTED ON AN AIRPLANE. Robert G.
Chilton. November 1954. 22p. diagrs., photo.,
tab. (NACA TN 3313)

The power spectrum of high-frequency turbulence in
the atmosphere was calculated from measurements
made in flight. The spectrum was found to display
an inverse variation with the square of frequency.
This variation is in agreement with the high-
frequency asymptote of the spectrum form commonly
associated with isotropic turbulence. Flow-direction
vanes were used to measure vertical and horizontal
components of gust velocity relative to the airplane
and normal to the flight direction. The power
spectral densities of the two components were, for
practical purposes, equal.

NACA TN 3315

TENSILE AND COMPRESSIVE STRESS-STRAIN
PROPERTIES OF SOME HIGH-STRENGTH SHEET
ALLOYS AT ELEVATED TEMPERATURES.
Philip J. Hughes, John E. Inge and Stanley B.
Prosser. November 1954. 32p. diagrs., photos.,
6 tabs. (NACA TN 3315)


A method is presented for determining the time-
dependent flow over a rectangular wring moving with a
supersonic forward speed and undergoing small
vertical distortions expressible as polynomials in-
volving spanwise and chordwise distances. Results
are expressed in terms of generalized indicial forces.
Numerical results for Mach numbers of 1.1 and 1.2
are given for polynomials of the first and fifth degree
in the chordwise and spanwise directions, respective-
ly, on a wing of aspect ratio 4.



NACA TN 3291

EXPERIMENTAL INVESTIGATION OF NOTCH-SIZE
EFFECTS ON ROTATING-BIEAM FATIGUE BE-
HAVIOR OF 75S-T6 ALUMINUM ALLOY. W. S.
Hyler, R. A. Lewis and H. J. Grover, Battelle
Memorial institute. Novem er 19154. 47lp. diagrs.,


This investigation was initiated to study the influ-
ence of size, particularly the notch size, on ex-
truded 7SS-T6 aluminum-alloy test specimens.
Unnotched and notched specimens with five different
minimum-section diameters were tested. For each
size a semicircular groove was tested and for the
largest diameter specimen a V-notch was also
tested. A method of surface preparation was se-
lected that would produce comparable surface
finishes in different-sized notched and unnotched
specimens.




NACA TN 3292

INFLUENCE OF EXPOSED AREA ON STRESS-
CORROSION CRACKING OF 248S ALUMINUM ALLOY.
William H. Corner and Howard T. Francis, Armour
Research Foundation. November 1954. 22p. diagrs.,
photos., tab. (NACA TN 3292)

Results are presented of a study of the "area effect"
in 24S aluminum alloy. This effect is the phenome-
non whereby small exposed areas show long times to
stress-corrosion failure, whereas large areas show
short times. The effects of stress level, degree of
sensitivity of the alloy, and hydrogen peroxide con-
centration in the corrosion medium were studied.
Hydrogen peroxide decomposition and the substitution
of oxygen for peroxide were also investigated.





NACA TN 3305

SOME MEASUREMENTS AND POWER SPECTRA OF
RUNWAY ROUGHNESS. James H. Walls, John C.
Houbolt and Harry Press. November 1954. 27p*
diagrs., tab. (NACA TN 3305)

Measurements of actual runway roughness obtained
by a profile-survey method (engineer's level) are
presented. Data were obtained from a survey of a
relatively rough runway and a smooth runway. The
results of this study are presented as roughness
profiles of the runways surveyed and in the form of
power spectra.







NACA
RESEARCH ABSTRACTS NO. 74

Results of tensile and compressive stress-strain
tests at temperatures up to 1,2000 F are presented
for SAE 4340, Hy-Tuf, Stainless W, and Inconel X
sheet materials which had ultimate tensile strengths
at room temperature in the 170 to 220 ksi range.
Representative tensile and compressive stress-strain
curves are given for each material at the test
temperatures. Secant and tangent moduli, obtained
from the compressive data, are included.


NACA TN 3343

SUBSONIC EDGES INr THIN-WING AND SLENDER-
BODiYTHEORY. Milton D. Van Dyke. November
1954. 26p. diagrs. (NACA TN 3343)






compressible flow, three-dimensional wings, sharp
edges, and slender bodies of revolution.


NACA TN 3345

ARRANGEMENT OF FUSIFORM BODIES TO RE-
DUCE THE WAVE DRAG AT SUPERSONIC SPEEDS.
Morris D. Friedman and Doris Cohen. November
1954 23p. diagrs. (NACATN 3345. Formerly
RM A51120)

Using linearized slender-body theory and reverse-
flow theorems, the wave drag of a system of fusi-
form, bodies at zero angle of attack and supersonic
speeds is studied to determine the effect of varying
the relative location of the component parts. It is
found that in certain arrangements the interference
effects are beneficial, and may even result in a two-
or three-body system having no more wave drag than
the principal body alone. The most favorable loca-
tion appears to be one in which the maximum cross
section of the auxiliary body is slightly forward of
the Mach cone! from the tail of the main body.




BIRITISH REPORTS




N-33504*

Forest Products Research Lab. (Gt. Brit.)
TRIAL OF TIMBERS FOR PLYWOOD MANUFAC-
TURE. SEPETIR PSEUDOSINDORA PALUSTRIS -
SARAWrAK. ( POUNS PER CUBIC FOOT AT
15 PER CENT MOISTURE CONTENT) PROGRESS
REPORT TWENTY-F;OUR. August 1954. 17p.
(Forest Products Research Lab.)

Thirteen 5 -ft 3 -in. billets were used for obtaining yield
data and two were used for processing trials. Good
green veneer was smooth and free from defects
except end splits, which were rather numerous and
hindered spooling. The extreme straightness of the
grain resulted in easy breakage and made high-speed
power spooling impossible. Generally the veneer


was similar in texture to afara. The heartwood and
sapwood had similar drying properties but the sap-
wood is a greenish color very distinct from the
pinkish heartwood. It is classified as being unsuited
to British mills but is technically sound for plywood
making and might be of value for this purpose in the
country of origin.


N-34046*

Royal Aircraft Establishment (Glt.Brit.)
THE EFFECT OF SURFACE FINIH ON THE
FATIGUE RESISTANCE OF TWO ALUIMINIUM
ALLOYS. N. J.F. Gunn. March 1954. 19p.
diagrs., 13 tabs. (RAE Tech. Note Met. 196)




pol or ished. Thehe fatigue~liae srantios frbTh


machined 0.30; fine machined 0.30; circumfer-
entially polished 0.32; longitudinally polished -
0.34.


N-34048*

Royal Aircraft Establishment (Gt. Brit.)
A GrENERALISATION OF THE NYQUIST STABILITY
CRITERION WITH PARTICULAR REFERENCE TO
PHASING ERROR. R. H. Merson. May 1954. 9p.
diagr. (RAE Tech.Note? GW 316)

The effect of phasing error on the stability of a two-
dimensional linear servomechanism is considered
and it is shown that the system will be stable if the
phase margin at the cut-off frequency exceeds the
phasing error. The more general. case of a number
of identical servos with cross coupling is investigated
and a generlization of the Nyquist criterion for
stability is formulated.



N-34049*

Royal Aircraft Establishment (Gt. Brit.)
THE DETERMINATION OF CALCIUM IN TITANIUM
METAL. H. J. Allsopp. May 1954. 9p., 7 tabs.
(RAE Tech. Note Met. 198)

In the method described, titanium is removed by
volatilization as the tetrachloride and the calcium are
estimated volumetrically using disodium-
ethylenediaminetetr a-ac estate. The effect of reagent
concentrations and some interfering elements has
been studied.



N-34051*

Royal Aircraft Establishment (Gt.Brit.)
DIELECTRIC MEASUREMENTS ON SOME LAMIN-
ATING RESINS AND THE EFFECT OF MOISTURE
ABSORPTION AND TEMPERATURE. A. A. Fyall
and J. H.Sewell. June 1954. 29p. diagrs.,
16 tabs. (RAE: Tech.Note Chem. 1233)










Values are given for the dielectric constant and
loss tangent of several resins having potential
application in the fabrication of glass-fiber rein-
forced radomes. Dielectric properties were meas-
ured on dry materials and also on materials sub-
jected, for varying periods of time, to tropical con-
ditioning in a humidity chamber. When the resin
composition could be varied, that giving the optimum
dielectric performance was ascertained. Where
possible, measured values were compared with those
obtained theoretically. The merits of the resins are
compared and their best values summarized.



N-34053*

Royal Aircraft Establishment (Gt.Brit.)
THE ESTIMATION OF SIZE ON GLASS FABRICS:
TWO NEW METHODS. E. Haythornthwaite and
R. B. King. June 1954. 15p. diagrs., photo.,
3 tabs. (RAE Tech. Note Chem. 1231)

Two new methods for the estimation of the size
content of glass fabrics have been examined. One
method depends on the gasometric estimation of
carbon dioxide formed from the carbon present on a
fabric sample. The other is a modified volatile loss
method suitable for control purposes and designed to
minimize the errors inherent in previous methods.
Excellent correlation between the results from the
two methods was obtained.




N-34055*

Royal Aircraft. Establishment (GL Brit.)
A NEW HIGH SPEED PHOTOGRAPHIC TECHNIQUE
APPLIED TO THE INVESTIGATION OF BUBBLES
BURSTING AT AN AIR-WATER INTERFACE.
R. L.Aspden. July 1954. 26p. diagrs., photos.'
3 tabs. (R AE Tech. Note Instn. 141)

The Colourflash system, a method of picture
separation by frequency discrimination whereby a

oultycn resculr d atnart wh h sims eentially
unlimited ,d edrthieupmn ic niti no g pmoblis

""ts einso desried in Aes apI hton tofd ih on e

of smoke filled bubbles is included,




N-34056'

Royal Aircraft Establishment (Gt.Brit.)
CALIBRATION AND DESCRIPTION OF THE EX-
HAUSTER SECTION OF THE HIGH ALTITUDE
TEST PLANT. E. Simpson and F. S. Margrie.
May 1954. 47p. diagrs., photos. (RAE Tech.Note
Aero 2303. Formerly RAE Tech. Memo. Aero 329)

This note gives a brief description of the Ex L. F. A.
Exhauster Plant now installed at R. A. E. as a unit
of the High Altitude Test Plant. The means used to
calibrate the plant are described. Results of cali-
brations indicate that with different combinations of
units, at mass flows varying from 1. 5 to 4. 5 lb/sec
pressures of 0.75 inch Hg to 2 inches Hg absolute


NACA
RESEARCH ABSTRACTS NO.74

could be obtained, and at mass flows varying from
3 lb/see to 8.5 lb/sec pressures of 1.35 inches Hg to
3.3 inches Hg were obtainable. Approximate rela-
tive efficiencies are given and lines of possible
improvement indicated. A detailed description of
the plant and the modifications introduced as a
result of operational experience is given. This note
is the first of a series describing and giving calibra-
tions of various sections of the High Altitude Plant.




N-34057*

Aeronautical Research Council (Gt. Brit.)
A SIMPLE APPROACH TO T`HE THEORY OF
SECONDARY FLOWS. J. H. Preston.
December 10, 1953. 19p. diagrs. (ARC 16, 394;
FM 1996; EA 320)

A simple method is developed for computing the
secondary trailing vorticity which arises when a
nonuniform stream is turned. The essential
results derived by Squire and Winter and by
Hawthorne are obtained with a minimum of analysis
and the reason for the secondary trailing vorticity
is physically evident. It is shown that for a sudden
deflection of a nonuniform stream, no trailing vor-
ticity is set up in the exit flow and hence there is
no secondary motion. It is also shown that for
small angles of deflection there is no net circulation
downstream of a cascade of finite dimensions and
it is inferred that this should be true also for large
deflections.



N-34106*

Royal Aircraft Establishment (Gt.Brit. )
A CROSSED-FIELDS MULTIPLIER. J. A. Roberts
and D. C.Pressey. April 1954. 88p. diagrs.,
photos., tab. (RAE Tech. Note Arm. 516)

Fed-back d-c amplifiers controlling crossed electric
and magnetic fields in a simple bucket cathode-ray
tbe VCRX. 3 0, e ale a naual pd otd lp ito be

do anaou c mue su pe f vlomect alesign,

dsrbed.b Direc neas rmentos owsothe stah He
the change of scale-factor with frequency does not
reach 1 percent until 10 kc/s. The power con-
sumpt-ion is only 50 w total. The zero drift is
small and the long-term stability excellent.





N-34111*

Aeronautical Research Council (Gt.Brit.)
ON THE TURBULENT BOUNDARY LAYER ON A
FLAT PLATE OF FINITE WIDTH. A. A. Townsend.
March 3, 1954. 10p. diagrs. (ARC 16,618;
FM 2042)

A theory is developed for edge effect on the turbu-
lent boundary layer on a flat plate of finite width by
considering the equations of motion for the flow.






NACA
RESEARCH ABSTRACTS NO. 74

N-34113*

Aeronautical Research Council (Gt. Brit.)
AN EXPERIMENTAL INVESTIGATION OF THE
INTERACTION OF A SHOCK WAVE WITH A SUB-
SONIC STREAM BOUNDED BY A WALL. D. W.
Holder, A. Chinneck and G. E. Gadd. February 11,
1954. 17p. diagrs., photos., tab. (ARC 16, 559,
FM 2025)

A study has been made of a model of a boundary
layer consisting of a subsonic stream bounded on
one side by a wall and on the other by a supersonic
main stream. The subsonic stream was approxi-
mately uniform at the point where it first met the
m~ain streaml, but it became progressively less
uniform downstream of this point because of mixing
with the main stream and the growth of the boundary
layer at the wall. Particular attention is given to
the pressure rise which takes place at the wall when
a wedge is attached to it. The pressure begins to
rise ahead of the apex of the wedge, this upstream
effect increases as the Mach number of the
secondary stream is reduced.



MISCELLANEOUS



NACA Rept. 1133

Errata No. 1 on MECHANISM OF START AND
DEVELOPMENT OF AIRCRAFT CRASH FIRES. "
I. `Irving Pinkel, G. Merritt Preston and Gerard J.
Pesman. 1954.


NACA TN 1315

Errata No. 1 on "FREE-FALLS AND PARACHUTE
DESCENTS IN THE STANDARD ATMOSPHERE. n
A. P. Webster. June 1947.


N-22389*

Advisory Group for Aeronautical Research and
Development. SOME ASPECTS OF COMBUSTION
OF LIQUID FUEL. Charles C. Graves and Melvin
Gerstein. (Scheveningen Netherlands Conference,
May 3-7, 1954) 26p. diagrs., photo. (AGARD
AG16/M10)


1osdee natems auc i dvd al poe s as

turbojet combustors as affected by fuel volatility,
spray characteristics, and the burning rate of
single drops is treated. The scope of this paper is
limited to a discussion of these factors normally
associated with diffusion flames.

N-25254*

Advisory Group for Aeronautical Research and
Development. A SCHEME OF AUTOMATIC DATA
REDUCTION FOR WIND TUNNELS. K. V. Diprose.
(London AGARD Conference, September 3-11, 1953)
17p. diagrs., photos. (AGARD AG9/Md5)


The need for automatic devices to speedup reduction
of data obtained from, wind tunnels in the RAE first
became urgent during the war. Up to the present
little of the analogue computing equipment suggested
in the original plan has been constructed. One com-
puter giving tunnel Mach number was constructed in
1948 and has been in operation ever since in the
10- by 'I-foot high-speed tunnel. A "breadboard"
model of a computer to calculate uncorrected force
coefficients has been made. The design of plotting
tables as an alternative form of output is well
advanced. Their principal use will be in intermit-
tent wind tunnels where time available for a complete
run does not permit printing typed answers.


N-30835A*

Advisory Group for Aeronautical Research and
Development. A NOTE ON THE UISE OF STRAIN
GAUGES IN WIND TUNNEL BALANCES. J. R.
Anderson. (London AGARD Conference,
September 3-11, 1953) 20p. diagrs., photo. (AGARD
AG10/M6)

This paper records some of the experience obtained
in the use of strain gage balances in the smaller
ehag-stpee dtindA tnels oidiAer odna sict rfr
little restricted in scope, being for the most part
limited to the use of bonded, resistance type elec-
trical strain gages in force and moment balances
intended for use with models ranging from about 0.4
in, to 1.5 in, in diameter. Much of the discussion
should prove of interest in the design of larger
installations: for example, the comments which are
made on the effects of temperature on the strain
gages.


N-33567*

Advisory Group for Aeronautical Research and
Development. FORMATION ET DEPOT DE
CARBON DANS LES FOYERS DE TURBO-
MACHINES D'AVIATION. C. Four$.
(Scheveningen Netherlands Conference, M~ay 3-7,
1954) 21p. diagrs., photos. (AGARD AG12/M8)

Operating with accepted fuels, modern reactionary
turbines are not seriously troubled by the formation
of carbon deposits. The change towards high
compression and towards the use of less powerful
fuels alter this situation. The role of certain
characteristics of fuels have been clarified. A




the walls.


N-33583*

DESIGN AND OPERATING TECHNIQUES OF VER-
TICAL SPIN TUNNELS. A. I. Neihouse. (Presented
to Wind TIunnel and Model Testing Panel of Advisory
Group for Aeronautical Research and Development,
Paris, France, November 2-6, 1954) 16p. diagrs.,
photos., 2 tabs.
Dynamic tests in a spin tunnel, aided by auxiliary
tools and techniques including analytical methods,
offers a practical method for a better understanding









of the spinning motion and the recovery therefrom.
Just how the moments and forces affect the picture>
and just what part the stability characteristics of the
airplane at large attitude angles play can be deter-
mined by proper research effort. That such effort
is necessary has been demonstrated in the past and
by experiences with current airplanes; that the
problem of spin recovery may become more critical
is likely. Adequate facilities to do spin research,
however, should lead to eventual solution of the
existing problem.



N-33609*

TRANSONIC WIND TUNNEL DEVELOPMENT OF
THE NATIONAL ADVISORY COMMITTEE FOR
AERONAUTICS. H. Julian Allen. (Presented to
Wind Tunnel and Model Testing Panel of Advisory
Group for Aeronautical Research and Development,
Paris, France, November 2-6, 1954) 22p. diagrs.,
photo.

The development of transonic wind tunnels at the
Langley and Ames Aeronautical Laboratories of the
NACA is discussed. Two types of wind tunnels
suitable for transonic wind tunnel operation are
described: the fixed geometry slotted-wall type and
the variable-geometry ventilated wall (slotted or
porous) type.


N-33610 *

DEVELOPMENT OF TWO HYPERSONIC TEST
FACILITIES AT THE NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS, AMES
AERONAUTICAL LABORATORY. H. Julian Allen.
(Presented to Wind Tunnel and Model Testing Panel
of Advisory Group for Aeronautical Research and
Development, Paris, France,November 2-6, 1954)
41p. diagrs., photos., tab.
Two facilities for hypersonic research which have
been developed at the Ame s Aeronautical Laboratory
are discussed. The first facility is a wind tunnel
which is of interest because of its ability to operate
continuously over a wide range of speeds with a not-
too-complex drive system; while the second is a
combination wind tunnel and ballistics range which
allows operation over an even wider range of speeds
up to speeds which are very difficult to attain in a
wind tunnel of the more conventional type.



UNPUBLISHED PAPERS



N-25802*

THE ROTATING WING AS A RADIATION PROBLEM.
(Der rotierende Tragfliigel als Strablungsproblem).
Wilhelm Ernsthausen. September 1954. 40p.
diagrs., photos. (Trans. from Zeitschrift fiir
angewandte Mathematik und Mechanik, v.31, no. 1/2,
Jan. /Feb., 1951, p. 20-35)

The forces exerted upon a wing rotating in a com-
pressible gas correspond to the reaction of its field
of sound. Therefore, they may be determined from


NACA
RESEARCH ABSTRACTS NO. 74

the energy capacity of this field. The investigation
is based on the idea of starting with the field of
sound and penetrating into the connections between
flow and wave concepts. It deals with the determi-
nation of the flow forces stemming from the energy
of the sound field which acts on the wing, with the
mass motion describing the lift finding its expres-
sion in the field in the immediate neighborhood, and
the radiation losses responsible for the drag finding
theirs in the field in the distance.



N-33628*

APPLICABILITY OF THE LAW OF SIMILIUDE TO
CAVITATION. (Zum Abhnlichkeitsgesetz fiir
Hohlraumbildungen). A. Reinhardt. April 1954.
34p. diagrs., photo. (Trans. from Forschung auf
dem Gebiete des Ingenieurwesens, Ed. B, v. 6,
p. 1-12, Jan. -Feb., 1935, Forschungsheft 370)

Conclusions, derived from the theory of similitude,
regarding aviation in flow engines were compiled,
supplemented, and investigated as to their practical
applicability. ?Two deviations from Newton's
similitude of flow were recognizable in the test
results. A number of tests were performed with a
Kaplan turbine in order to check theoretical results.




DECLA S SIFI ED NACA REPORTS







THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL, 10/29/54:

RM 8J50
RM 51D)23
RM E6L11
RM E7JO1
RM L8A14
RM: L50LO8
RM L52C31
RM LS2J14



NACA RM A7L24

HIGH-SPEED WIND-TUNNEL TESTS OF A 1/78-
SCALE MODEL OF THE LOCKHEED YP-80A AIR-
PLANE. Robert N. 01son and Leslie F. Lawrence.
May 28, 1948. 52p.diagrs., photos. (NACA RM
A7L24) (Declassified from Confidential, 10/12/54)

This report contains the results of a high-speed
wind-tunnel investigation of a 1/78-scale model of the
Lockheed YP-80A airplane, including a comparison
of the relative aerodynamic characteristics of the
1/78-scale model, a i1/-scale model, and a full-
scale YP-80A airplane. Also included are the longi-
tudinal stability and control characteristics of the
1/78-scale model with 00 and 450 leading-edget
sweepback of the horizontal and vertical tail surfaces.








NACA
RESEARCH ABSTRACTS NO. 74


NACA RM L51K29

PROPELLER LIFT AND THRUST DISTRIBUTION
FROM WAKE SURVEYS OF STAGNATION CONDI-
TIONS. Robert E. Davidson. January 1952. 19p*
diagrs. (NACA RM L51K29) (Declassified from
Confidential, 10/12/54)

This paper gives formulas for propeller lift and
thrust distribution in terms of wake-survey
measurements of stagnation pressure rise through
the propeller. Substantiative data are also
included.





NACA RM L51LO6

PROPELLER INDUCED ANGLES OF ATTACK AND
SECTION ANGLES OF ATTACK FOR THE NACA
10-(3)(066)-03, 10-(3)(049)-03, 10-(3)(090)-03,
10-(5)(066)-03, AND 10-(0)(066)-03 PROPELLERS.
William B. Igoe and Robert E5. Davidson. May 1952.
80p. diagrs., 10 tabs. (NACA RM L51LO6)
(Declassified from Confidential, 10/12/54)

This paper presents the results of an induced angle-
of-attack calculation using a method applicable to a
propeller with arbitrary circulation distribution.
Tables of induced angles of attack and section angles
of attack and curves of wake-survey results are pre-
sented for the NACA 10-(3)(066)-03, 10-(3)(049)-03,
10-(3)(090)-03, 10-(5)(066)-03, and 10-(0)(066)-03
propellers. A brief description of the method of cal-
culating propeller induced angles of attack is given.




NACA RM L51L28

THE EFFECT OF BLADE-SECTION CAMBER ON
THE STATIC CHARACTERISTICS OF THREE NACA
PROPELLERS. John H. Wood and John M. Swibart.
April 1952. 40p. diagrs., photos. (NACA RM
L51L28) (Declassified from Confidential, 10 12/54)

This paper contains the effect of blade-section
camber on the static characteristics of the NACA
10-(0)(066)-03, 10-(3)(066)-03, and 10-(5)(066)-03
propellers (design lift coefficients of 0, 0.3, and 0.5,
respectively). Blade angles from Go to 160 were
tested over a tip Mach number range from. 0.28 to
1.02. The results indicate that the use of camber in
the propeller design offers advantages in propeller
performance and increases in stall-flutter speed at
zero advance.



NACA RM L52AO3

HINGE-MOMENT AND OTHER AERODYNAMIC
CHARACTERISTICS AT TRANSONIC SPEEDS OF A
QUARTER-SPAN SPOILER ON A TAPERED 450
SWEPTBACK WING OF ASPECT RATIO 3.
Joseph E. Fikes. February 1952. 22p. diagrs.,
photo. (NACA RM L52AO3) (Declassified from
Confidential, 10/12/54)


Lift, drag, pitching moment, rolling moment, yawing
moment, and spoiler hinge moment were obtained at
transonic speeds by testing in the high velocity field
over a reflection plane on the side wall of the Langley
high-speed 7- by 10-foot tunnel on a wing having a
quarter-chord sweepback of 45.58o, an aspect ratio of
3, a taper ratio of 0.5, and an NACA 64A010 section
employing an inboard quarter-span plug spoiler. The
investigation was made over a limited projection and
angle-of-attack range from a Mach number of 0.70 to
1.10.



NACA RM L52A09

EFFECT OF CURRENT DESIGN TRENDS ON AIR-
PLANE SPINS AND RECOVERIES. Anshal I.
Neihouse. January 1952. 6 diagrs. (NACA RIM
LS2A09) (Declassified from Confidential, 10/ 12/54)

Charts are presented which indicate that, owing to
design of airplanes for transonic and supersonic
flight, a wide range of mass distribution of the air-
plane is possible with accompanying effects upon the
nature of the spin and upon the requirements for
recovery.



NAC"A RM L52A~l

PRELIMINARY INVESTIGATION OF CONTROL
CHARACTERISTICS AT TRANSONIC SPEEDS OF A
TAPERED 450 SWEPTBACK WING OF ASPECT
RATIO 3 HAVING A HORN-BALA~NCED FULL-SPAN
CONTROL. John G. Lowry and Joseph E. Fikes.
April 1952. 22p. diagrs., photo. (NACA RM
L52All) (Declassified from Confidential, 10/12 541

An experimental investigation was made at transonic
speeds by testing in the high-velocity field over a
side~wall reflection plane to determine hinge-moment
and effectiveness characteristics of a horn-balanced
control on an aspect-ratio 3, 450 sweptback wing.
Lift, rolling moments, and control hinge moments
were obtained through a limited angle-of -attack and
control-deflection range from Mach numbers 0.70 to
1.10.



NACA RM L52A14

PRELIMINARY INVESTIGATION OF THE DRAG
CHARACTERISTICS OF THE NACA RM-10 MISSILE
AT MACH NUMBERS OF 1.40 AND 1.59 IN THE
LANGLEY 4- BY 4-FOOT SUPERSONIC TUNNEL.
Lowell E. Hasel, Archibald R. Sinclair and Clyde V.
Hamilton, April 1952. 49p. diagrs., photos, 3 tabs.
(NACA RM L52A14) (Declassified from Confidential,
10 12 54)1

A 0.287-scale model of the RM-10 has been tested in
the Langley 4- by 4-foot supersonic tunnel at Mach
numbers of 1.59 and 1.40. Drag data were obtained
from a conventional sting-supported model and from
a wire-supported model. Base drag, skin-friction
drag, forebody pressure drag, fin drag, and fin inter-
ference drag are discussed. The data at a Mach
number of 1.59 are compared with that from other
facilities.










NACA RM L52A17

ERROR IN AIRSPEED MEASUREMENT DUE TO
STATIC-PRESSURE FIELD AHEAD OF AN OPEN-
NOSE AIR-INLET MODEL AT TRANSONIC SPEEDS.
Thomas C. O'Bryan. March 1952. 17p. diagrs.,
photos. (NACA RM L52A17) (Declassified from
Confidential, 10/12/54)

Contains measurements of the static-pressure coef-
ficient at several distances ahead of an open-nose
air-inlet body at zero angle of attack, in the tran-
sonic speed range, by the NACA wing-flow method*
Measurements are presented to show the effect of
inlet-velocity ratio on the static pressure ahead of
the inlet. Comparison of the experimental variation
of the static-pressure coefficient ahead of the inlet
is made with incompressible theory. Calculations
are presented to show the effect of changing inlet
geometry on the static-pressure coefficient ahead of
the inlets.






NACA RM L52A22

THE CALCULATION OF CERTAIN STATIC AERO-
ELAS*TIC PHENOMENA OF WINGS WITH TIP
TANKS OR BOOM-MOUNTED LIFTING SURFACES.
Franklin W. Diederich and Kenneth A. Foss.
August 1952. 55p. diagrs., 2 tabs. (NACA
RM L52A22) (Declassified from Confidential,
10/12/54)

A method is presented for calculating static aero-
elastic phenomena of wings with concentrated aero-
dynamic forces at the tip. Some static aeroelastic
characteristics of an unswept wing with a tip tank and
of a sweptback wing with several boom-mounted lift-
ing surface configurations have been calculated.
These calculations indicate that such a lifting.surface
merits consideration as a device for relieving ad-
verse aeroelastic effects.





NACA RM L52A23

PRELIMINARY INVESTIGATION AT TRANSONIC
SPEEDS OF THE EFFECT OF BALANCING TABS
ON THE HINGE-MOMENT AND OTHER AERODY-
NAMIC CHARACTERISTICS OF A FULL-SPAN FLAP
ON A TAPERED 450 SWEPTBACK WING OF ASPECT
RATIO 3. Vernard E. Lockwood and Joseph E.
Fikes. April 1952. 27p. diagrs., photo., (NACA
RM L52A23) (Declassified from Confidential,
10/12/54)

An experimental investigation was performed at tran-
sonic speeds in the high-velocity-flow field over a
reflection plane in the Langley high-speed 7- by 10-
foot tunnel to determine the balancing characteristics
of an inset, an attached, and a detached tab on an
aspect-ratio 3, 450 sweptback wing. Lift, rolling
moments, and flap hinge moments were obtained
through a range of angles of attack from Go to 160
and a Mach number range from 0.7 to 1.1. Data are
presented as parameters of the flap and tab.


NACA
RESEARCH ABSTRACTS NO.74


NACA RM L52A25

LOW-SPEED STABILITY CHARACTERISTICS OF A
COMPLETE MODEL WITH A WING OF W PLAN
FORM. Edward C. Polhamus and Robert E. Becht.
April 1952. 27p. diagrs., photo., tab. (NACA
RM L52A25) (Declassified from Confidential,
10/12/54)

An investigation was conducted to determine the low-
speed static longitudinal and lateral stability charac-
teristics of a complete model equipped with a W wing
of aspect ratio 6, taper ratio 0.6, and panel sweep
angles of 450. Included in the investigation are ex-
ploratory tests made in an attempt to delay the early
separation in the wing-panel junctures by use of
fences, vortex generators, and chord extensions.



NACA RM L52B11

INVESTIGATION OF THE HYDRODYNAMIC STA-
BILITY AND RESISTANCE OF TWO STREAMLINE
FUSIELAGES. Bernard Weinflash and Charles L.
Shuford, Jr, April 1952. 32p.diagrs., photos.,
tab. (NACA RM L52B11) (Declassified from
Confidential, 10/12/54)

An investigation of a dynamic model was made to
determine the effects of hull form, gross load, and
aerodynamic trimming moments on the trim limits,
trim, hydrodynamic moment, hydrodynamic resist-
ance, total resistance, and rise of two streamline
fuselage configurations modified by chine strips.




NACA RM L52B13

EFFECTS OF SOME PRIMARY VARIABLES OF
RECTANGULAR VORTEX GENERATORS ON THE
STATIC-PRESSURE RISE THROUGH A SHORT
DIFFUSER. E. Floyd Valentine and Raymond B.
Carroll. May 1952. 32p. diagrs., photo., tab.
(NACA RM L52B13) (Declassified from Confidential,
10/12/54)

An investigation was made of a 2:1 area ratio dif-
fuser of length equal to the inlet diameter with sepa-
rate variation of several basic parameters for sim-
ple nontwisted rectangular vortex generators over a
considerable range of inlet-boundary-layer thickness.
Optimum values from the standpoint of static pres-
sure rise were determined for angle of attack,
spacing, aspect ratio, and span-to-inlet boundary-
layer thickness. The static-pressure rise obtainable
by conforming to these findings is given over the
range of inlet-boundary-layer thickness.



NACA RM L52B15a

AN INVESTIGATION OF LONGITUDINAL CONTROL
CHARACTERISTICS OF A WING-TIP CONTROL
SURFACE ON A SWEPTBACK WING AT TRANSONIC
SPEEDS BY THE NACA WING-FLOW METHOD.
James P. Trant, Jr. June 1952. 23p. diagrs.,
photos., tab. (NACA RM L52B15a) (Declassified
from Confidential, 10/12/54)







NACA
RESEARCH ABSTRACTS NO. 74

Longitudinal control effectiveness of a full-chord
wing-tip control on a wing having 350 sweepback 12
percent thickness perpendicular to the quarter-chord
line, an aspect ratio of 3.01, and a taper ratio of
0.605 was determined by the NACA wing-flow method
at Mach nubers from about 0.65 to 1.1. Meatsure-
ments were made of normal force, pitching moment,
hinge moment, and angle of attack for three control
deflections. One control deflection was also tested
with an end plate attached to the root chord of the
control surface.



NACA RM L52Bl8

THE EFFECTS ON THE AERODYNAMIC CHARAC-
TERISTICS OF VARYING THE WING THICKNESS
RATIO OF A TRIANGULAR WING-BODY CONFIG-
URATION AT TRANSONIC SPEEDS FROM TESTS BY
THE NACA WING-FLOW METHOD. Albert W. Hall
and James M. McKay. Apri 1952. 27p. diagrs.,
photo., 2 tabs. (NACA RM L52B18) (Declassified
from Confidential, 10/12/54)

Tests were made by the NACA wing-flow method at
Mlach numbers from 0.75 to 1.07 to determine the
aerodynamic characteristics of three triangular wing-
fuselage models which differed only in wing thickness-
chord ratio. All three wings had an aspect ratio of
2.31 with 6-, 9-, and 12-percent-thick biconvex
sections and the fuselage had a fineness ratio of 12.
Measurements were made of normal force, chord
force, and pitching moment for various angles of
attack. The test Reynolds number was 1.5 x 106.



NACA RM L52B25

EFFECTS OF HORIZONTAL-TAIL POSITION AND
ASPECT RATIO ON LOW-SPEED STATIC LONG~
TUD)INAL STABILITY AND CONTROL CHARACTER-
ISTICS OF A 600 TRIANGULAR-WING MODEL HAV-
ING TWIN TRIANGULAR ALL-MOVABLE TAILS.
Byron M. Jaquet. May 1952. 45bp. diagrs., photos.
(NACA RM LS2B25) (Declassified from Confidential,
10/12/54)

Presents results of a low-speed investigation made
in Langley stability tunnel to determine effects of
horizontal tail position and aspect ratio on low-speed
static longitudinal stability and control characteris-
ties of a 600 triangular-wing model having twin-
triangular-all-movable tails. All the force tests
were made at a Mach nuber of 0.17 and a Reynolds
number of 2.06 x 106. Comparison of results for
single and twin tails are made.



NACA RM LS2B26

FLIGHT MEASUREMENTS OF THE EFFECTS OF
SURFAE CONDITION ON THE SUPERSONIC DRAG
OF FIN-STAB.IZ~ED PARABOLIC BODIES OF REV-
OLUTION. H. Herbert Jackson. Mnay 1952. 17p.
diagrs., photos. (NACA RM L52B26) (Declassified
from Confidential, 10/12/54)

Some measurements of the effects of various types of
surface roughness on total drag and base drag were
obtained on bodies of revolution in free flight. The


Mach number and Reynolds number ranges varied
from 0.80 to 2.2 and 15 x 106 to 75 x 106, respec-
tively. The experimental drag coefficients of models
having wavy surfaces, sand-coated surfaces, and
pitted surfaces are compared ~with the drag coeffi-
cients of smooth models of the same configurations.



NACA RM L52C24

SKIN-FRICTION DRAG AND BOUNDARY-LAYER
TRANSITION ON A PARABOLIC BODY OF REVO-
LUTION (NACA RM-10) AT A MACH NUMBER OF
1.6 IN THE LANGLEY 4- BY 4-FOOT SUPERSONIC
PRESSURE TUNNEL. K. R. Czarnecki and Jack E.
Marte. May 1952. 24p. diagrs., photos. (NACA
RM L52C24) (Declassified from Confidential,
10/12/54)

There are presented the results of an investigation
at M = 1.6 and over a Reynolds number range from
2 x 106 to 40 x 106 of the skin-friction drag and
boundary-layer transition of a body of revolution
(NACA RM-10). The results are compared with
other available experimental data and with theory.



NACA RM L52C27

AN APPLICATION OF THE ROCKET-PROPELLED-
MODEL TECHNIQUE TO THE INVESTIGATION OF
LOW-LIFT: BUFFETING AND THE RESULTS OF
PRELIMINARILY TESTS. Homer P. Mason and
William N. Gardner. September 1952. 19p.
diagrs., photos. (NACA RM LS2C27) (Declassified
from Confidential, 10/12/54)

An application of the rocket-propelled-model tech-
nique for the investigation of low-lift buffeting is pre-
sented, together with data obtained from three pre-
liminary tests by this technique. A correlation be-
tween low-lift buffeting, wing dropping, and changes
of trim normal force is shown.



NACA RM L52D08a

SMALL-SCALE TRANSONIC INVESTIGATION OF
THE EFFECTS OF PARTIAL-SPAN LEADING-EDGE
CAMBER ON THE AERODYNAMIC CHAACTERIS-
TICS OF A 500 3g* SWEPTBACK WING OF ASPECT
RATIO 2.98. William J. Alford, Jr. and Andrew L.
Byrnes, Jr. June 1952. 28p. diagrs., photo., tab.
(NACA RM L52D)08a) (Declassified from Confiden-
tial, 10/12/54)

An investigation of two semispan wings having the
same plan form wras made in the Langley high-speed
7- by 10-foot tunnel over a Mach number range from
0.70 to 1.10 and a mean-test Reynolds nuber range
from 745,000 to 845,000 to determine the eff~ets of
partial-span leading-edge camber on the aerodynamic
characteristics of a sweptback wing. This paper pre-
sents the results of the wing-alone and wing-fuselage
configurations with the wing quarter-chord line swept-
back 500 38' taper ratio 0.45, and NACA 64rA-series
airfoil sections tapered in thickness ratio. Lift,
drag, pitching moment, and root-bending moment
were obtained for all configurations.











NACA RM LS2D15

A PRELIMINARY INVESTIGATION OF THE STATIC
AND DYNAMIC LONGITUDINAL STABILITY OF A
GRUMBERG HYDROFOIL SYSTEM. Norman S.
Land, Derrill B. Chambliss and William W. Petynia.
September 1952. 48p. diagrs., photos. (NACA
RM L52D15) (Declassified from Confidential,
10/12/54)

A preliminary investigation was made in order to de-
termine the static and dynamic longitudinal stability
characteristics and the force characteristics of the
Grunberg hydrofoil system, comprising a main lifting
hydrofoil and planing-surface stabilizers. The static
stability of the system was determined for several
locations of the center of gravity. In smooth water
the response to a sudden distrubance was observed.
The behavior in waves of various lengths was ob-
served at several speeds, moments of inertia, and
locations of the center of gravity. The effects of
gross load, hydrofoil and stabilizer incidence,
speed, and center-of-gravity location on the lift-drag
ratio were determined.




NACA RM L52D21

EFFECTS OF PLAN FORM, AIRFORL SECTION,
AND ANGLE OF ATTACK ON THE PRESSURES
ALONG THE BASE OF BLUNT-TRAILING-EDGE
WINGS AT MACH NUMBERS OF 1.41, 1.62, AND
1.96. Kennith L. Goin. September 1952. 52p.
photos., diagrs. (NACA RM L52D21) (Declassified
from Confidential, 10/12/54)

Base pressures were measured at angles of attack of
00 to 150 on two groups of untapered wings of aspect
ratio 2.7, the first group being unswept and the sec-
ond group having 450 of sweepback. Each group in-
cluded airfoil sections with maximum thickness
ratios of 3 to 10 percent and with varying amounts of
trailing-edge bluntness. Also included in the inves-
tigation to indicate additional effects of wing plan
form twr ea 5450 delta wing and a rectangular wng of

transition at Reynolds numbers of between 1 x 106
and 2 x 106.


NACA RM L52D23a

A THEORETICAL AND EXPERIMENTAL INVESTI-
GATION OF THE LIFT AND DRAG CHARACTERIS-
TICS OF A HYDROFOIL AT SUBCRITICAL AND
SUPERCRITICAL SPEEDS. Kenneth L. Wadlin,
Charles L. Shuford, Jr. and John R. McGehee. July
1952. 53p. diagrs., photo., tab. (NACA
RM L52D23a) (Declassified from Confidential,
10/12/54)

A theoretical and experimental investigation was
made of the effect of the free-water surface and rigid
boundaries on the lift and drag of an aspect-ratio-10
hydrofoil at both subcritical and supercritical speeds.
Approximate theoretical solutions for the effect of
the free-water surface and rigid boundaries on lift
and drag at supercritical speeds were developed. An
approximate theoretical solution for the effect of
these boundaries on drag at subcritical speeds was
also presented. The agreement between theory and


NACA
RESEARCH ABSTRACTS NO.74

experiment at both supercritical and subcritical
speeds is satisfactory for engineering calculations of
hydrofoil characteristics from, aerodynamic data.



NACA RM L52D25

CONTROL CHARACTERISTICS AT TRANSONIC
SPEEDS OF A LINKED FLAP AND SPOILER ON A
TAPERED 450 SWEPTBACK WING OF ASPECT
RATIO 3. Vernard E. Lockwood and Joseph E.
Fikes. July 1952. 24p. diagrs., photo. (NACA
RM L52D25) (Declassified from Confidential,
10/12/54)

Lift, drag, pitching-moment, rolling-moment,
yawing-moment, and hinge-moment characteristics
of the model were obtained at transonic speeds by
testing in the high-velocity field over a reflection
plane on the side wall of the Langley high-speed 7- by
10-foot tunnel. The wing had a quarter-chord line
sweepback of 45.580, an aspect ratio of 3, a taper
ratio of 0.5, and was approximately 7.6 percent thick.
A quarter-span plug-type spoiler which was linked to
rotate with the flap in the same direction was employ-
ed as a control-balancing device. The investigation
was made over a limited projection and angle-of-
attack range from a Mach number of 0.70 to 1.10.





NACA RM L52D30

PRESSURE DISTRIBUTIONS ON BODIES OF REVO-
LUTION AT SUBSONIC AND TRANSONIC SPEEDS.
Richard I. Cole. July 1952. 47p. diagrs., photos.,
tab. (NACA RM L52D30) (Declassified from Confi-
dential, 10/12/54)

Pressure distributions along the slender body and
along prolate spheroids of fineness ratio 3 to 20 are
compared at subsonic and transonic speeds with
estimates at angles of attack up to 20 The com-

aon sh oie a be p eiced wth fi uracy.




NACA RM L52E12

SMALL-SCALE TRANSONIC INVESTIGATION OF
THIE EFFECTS OF FULL;-SPAN AND PARTIAL-
SPAN LEADING-EDGE FLAPS ON THE AERO-
DYNAMIC CHARACTERISTICS OF A 500 38' SWEPT-
BACK WING OF; ASPECT RATIO 2.98. Kenneth P.
Spreemann and William J. Alford, Jr. July 1952.
31p. diagrs., photo. (NACA RM L52E12) (De~classi-
fied from Confidential, 10/12/54)

An investigation was made in the Langley high-speed
7- by 10-foot tunnel over a Mach number range of
0.70 to 1.10 to determine the effects of a number of
full-span and partial-span leading-edge-flap deflec-
tions on the aerodynamic characteristics of a semi-
span model with the quarter-chord line sweptback 500
38', aspect ratio 2.98, taper ratio 0.45, and NACA
64A series airfoil sections tapered in thickness ratio.
Lift, drag, pitching moment, and bending moment
were obtained for all configurations.







NACA
RESEARCH ABSTRACTS NO.76


NACA RM L52E19

A THEORETICAL INVESTIGATION OF THE
EFFECT OF A TARGET SEEKER SENSITIVE TO
PITCH ATTITUDE ON THE DYNAMIC STABILITY
AND RESPONSE CHARACTERISTICS OF A SUPER-
SONIC CANARD MISSILE, CONFIGURATION.
Ordway B. Gates, Jr. and Albert A. Schy. August
1952. S4p. diagrs., photo., tab. (NACA
RM LS2E19) (D~eclassified from Confidential,
10/12/54)

A theoretical investigation is presented of the longi-
tudinal dynamic characteristics of an automatically
stabilized supersonic canard missile equipped with a
target seeker sensitive to pitch attitude. Effects of
seeker gain, time delay, and nonlinearities, which
include various dead spots in the target seeker, are
considered. The results indicate that satisfactory
stability and response characteristics may be ob-
tained for the configuration considered subsequent to
command or regulatory inputs but a dead spot in the
target seeker causes undesirable steady-state
errors in pitch attitude.



NACA RM L52Fl6

LATERAL-CONTROL INVESTIGATION AT TRAN-
SONIC SPEEDS OF RETRACTABLE SPOILER AND
PLUG-TYPE SPOILER-SLOT AILERONS ON A
TAPERED 600 SWEPTBACK WING OF ASPECT
RATIO 2. TRANSONIC-BUMP METHlOD. Alexander
D. Hammond and James M. Watson. August 1952.
19p. diagrs. (NACA RM LS2Fl6) (Declassified
from Confidential, 10/12/54)

Results and discussion of a lateral-control inves-
tigation of retractable spoiler and plug-type spoiler-
slot ailerons on a wing with 600 sweepback of the
quarter-chord line, aspect ratio 2, taper ratio 0.6,
and an NACA 65A006 airfoil section are presented.
Data were obtained through an angle-of-attack range
of -2o to 180, a projection range of -0.056 to 0.045
chord, and a Mach number range of 0.619 to 1.167.
The plug-type spoiler-slot aileron was more ef.
fective in producing roll than the retractable spoiler
aileron, particularly at the high angles of attack
below the speed of sound.






NACA RM L52F24

CALIBRATION OF A COMBINED PITOT-STATIC
TUBE AND VANE-TYPE FLOW1 ANGULARITY IN-
DICATOR AT TRANSONIC SPEEDS AND AT LARGE
ANGLES OF ATTACK OR YAW. Albin O. Pearson
and Harold A. Brown. September 1952. 25p.
diagrs., photos. (NACA RM L52F24)j (Declassified
from Confidential, 10/12 54)

The calibration of a combination pitot-static tube
and vane-type flow-angularity indicator is presented
for a Mach number range of approximately 0.60 to
1.11, an angle-of-attack range of -100 to 250, and
an angle-of-yaw range of -200 to 20o.


NACA RM L52GO3

THE EFFECT OF VARIOUS AERODYNAMIC BAL-
ANCES ON THIE LOW-SPEED LATERAL-CONTROL
AND HINGE-MOMENT CHARACTERISTICS OF A
0.20-CHORD PARTIAL-SPAN OUTBOARD AILERON
ON A WING WITH LEADING EDGE SWEPT BACK
51.30. Alexander D. Hammond. September 1952.
40 p. diagrs., photo., tab. (NACA RiM L52GO3)
(Declassified from Confidential, 10 12 54)1

A wind-tunnel investigation at low speeds was made
to determine the lateral-control and hinge-moment
characteristics of an unsealed, plain-radius-nose,
20t-percent-chord, flat-sided, partial-span outboard
aileron equipped with either an overhang, a paddle,
or a spoiler balance on a wing with leading edge
swept back 51.30, aspect ratio of 3.05, taper ratio
of 0.49, and NACA 651 -012 airfoil sections perpendi-
cular to the 55.6-perc'ent chord line.



NACA RM L52G08

HINGE-MOMENT AND CONTROL-EFFECTIVENESS
CHARACTERISTICS OF AN OUTBOARD FLAP WITH
AN OVERHANG NOSE BALANCE ON A TAPERED
350 SWEPTBACK WING OF ASPECT RATIO 4.
TRANSONIC-BUM~P METHOD. Robert F. Thompson
and William C. Moseley, Jr. August 1952. 51p.
diagrs. (NACA RMI LS2G08) (Declassified from
Confidential, 10 12/54)

Lift, pitching-moment, rolling-moment, and flap
hinge-moment coefficients were obtained through the
transonic speed range on a wing having a quarter-
chord sweepback of 350, an aspect ratio 4, a taper
ratio of 0.6, and an NACA 65A006 airfoil, section
parallel to free stream. The wing had a 30-percent-
chord 0.43-span outboard flap-type control with a
22-percent-flap-chord overhang nose balance. The
investigation was made through an angle-of-attack
range of -6o to 160, a Mach number range from 0.6
to 1.10, and a range of flap deflections which varied
from about f240 at a Mach number of 0.60 to +120
at a Mach number of 1.10.




NACA RM L52G10

LOW-SPEED LATERAL-CONTROL INVESTIGATION
OF A FLAP-TYPE SPOILER AILERON WITH AND
WITHOUT A DEFLECTOR AND SLOT ON A 6-
PERCENT-THICK, TAPERED, 450 SWEPTBACK
WIG OF ASPECT RATIO 4. James M. Watson.
September 1952. 11p. diagrs. (NACA RM L52GIDI
(Declassified from Confidential, 10/12/54)

Results are presented of an investigation to deter-
mine the lateral-control characteristics of a deflec-
tor and slot arrangement in conjunction with a flap-
type spoiler aileron on a 6-percent-thick wing with
450 sweepback of the quarter-chord line, an aspect
ratio of 4, and a taper ratio of 0.6. Data were ob-
tained through a range of angle of attack from -260
to 600, projections of 0 and 0.05 chord for both the
flap-type spoiler aileron and the deflector. The
spoiler-deflector-slot combination was more effee-
tive than the spoiler alone in producing rolling
moment.







NACA
RESEARCH ABSTRACTS NO.74

The intensity of reflections from walls of shock
waves, which were created by a bullet traveling near
the speed of sound, was observed by shadow and
schlieren methods. Results indicated that reflected
waves are greatly reduced by the use of a wall ma-
terial having good sound-absorbing characteristics,
or by specially designed wall configurations.


NACA RM L52G31a

EFFECT OF THICKNESS, CAMBER, AND THICK-
NESS DISTRIBUTION ON AIRFOIL CHARACTERIS-
TICS AT MACH NUMBERS UP TO 1.0. Bernard N.
Daley and Richard S. Dick. October 1952. 76p.
photos., diagrs., tab. (NACA RM L52G31a)
(Declassified from Confidential, 10/12/54)

A modified open-throat-type wind tunnel developed
for the purpose of obtaining two-dimensional -airfoil
data at Mach numbers near 1.0 is presented and dis-
cussed. Tests of a group of related NACA airfoil
sections varying in maximum thickness, design lift
coefficient, and thickness distribution have been con-
ducted in this wind tunnel at Mach numbers of 0.3 to
about 1.0, and at corresponding Reynolds numbers
from 0.7 x 106 to 1.6 x 106. Normal-force, drag,
and moment coefficients are presented, together with
representative schlieren photographs and pressure-
distribution diagrams.



NACA RM L52HO4

HEAT TRANSFER AND SKIN FRICTION FOR TUR-
BULENT BOUNDARY LAYERS ON HEATED OR
COOLED SURFACES AT HIGH SPEEDS, Coleman
duP. Donaldson. October 1952. 20p.diagr.,
3 tabs. (NACA RM L52HO4) (Declassified from
Confidential, 10/12/54)

The method presented in NACA TN 2692 for evaluat-
ing the skin friction of a turbulent boundary layer in
compressible flow on an insulated surface is extended
to evaluate the turbulent skin friction and heat trans-
fer in compressible flow on a surface which is heated
or cooled. The results of this analysis are in good
agreement with the heat transfers measured in flight
on the NACA RM-10 missile up to Mach number of
3.8.


NACA RM L52HO7

INVESTIGATION OF THREE TAPERED 450 SWEPT-
BACK CAMBERED AND TWISTED WINGS COVER-
ING A SIMULTANEOUS VARIATION IN ASPECT
RATIO AND THICKNESS RATIO AND OF ONE RE-
LATED SYMMETRICAL WING AT TRANSONIC
SPEEDS BY THE WING-FLOW METHOD. Harold I.
Johnson. March 1953. 55p. diagrs., photos., 2 tabs.
(NACA RM L52H07) (Declassified from Confidential,
10/12/54)

An investigation at transonic speeds was made to
study the effects of aspect ratio of a 450 sweptback
wing for the case where thickness ratio is varied in
a logical manner with aspect ratio. The effects of
camber and twist were also investigated. The Mach
numbers ranged from 0.65 to 1.17 and the Reynolds
numbers ranged from 230,000 to 620,000. Results
showed that the lowest-aspect-ratio, thinnest wing
tested had the best aerodynamic characteristics in


NACA RM L52G16

TESTS OF A CENTERING SPRING USED AS AN AR-
TIFICAL FEEL DEVICE ON THE ELEVATOR OF A
FIGHTER AIRPLANE. James J. Adams and James
B. Whitten. September 1952. 18p. diagrs., photo.,
tab. (NACA RM L52G16) (Declassified from
Confidential, 10/12/54)

Tests were made of a centering spring, which gave
no variation of force gradient with impact pressure
used as an artificial feel device for the elevator of a
Chance Vought F4U-4B airplane equipped with power
controls. The centering spring was unsatisfactory
because of the large stick force encountered in land-
ing. When a preloaded spring was included in the
system to relieve the large landing force, and a bob-
weight was added to increase the force per g, the
system satisfied the minimum handling-qualities re-
quirements, but was still considered to provide in-
sufficient centering tendency at high speeds,


NACA RM L52G22

THE EFFECTS OF CAMBER AND LEADING-EDGE-
FLAP DEFLECTION ON THE PRESSURE PULSA_
TIONS ON THIN RIGID AIRFOILS AT TRANSONIC
SPEEDS. Milton D. Humphreys and John D. Kent.
October 1952. 26p. diagrs., photos., tab. (NACA
RM L52G22) (Declassified from Confidential,
10/12/54)

The effect of camber and leading-edge-flap deflection
on the pressure pulsations on thin rigid airfoils at
Mach numbers from 0.5 to 1.0 has been investigated
for 6-percent-thick NACA 64A-series airfoils. The
results of the pressure-pulsation investigation on
these airfoils indicated that at the higher normal-
force coefficients, either camber or leading-edge-
flap deflection reduced the pressure-pulsation level
over the entire chord'




NACA RM L52G24

EXPERIMENTAL INVESTIGATION OF THE EF-
FECTS OF VORTEX GENERATORS ON THE MAXI-
MUM LIFT OF A 6-PERCENT-THICK SYMMETRI_
CAL CIRCULAR-ARC AIRFOIL SECTION. William
J. Bursnall. October 1952. 16p. diagrs., 2 tabs.
(NACA RM L52G24) (Declassified from
Confidential, 10/12/54)

The results of an experimental investigation of the
effectiveness of several configurations of vortex
generators in increasing the maximum lift of a 6.
percent-thick symmetrical circular-are airfoil sec-
tion indicated that none of the configurations em-
ployed substantially increased the maximum lift.



NACA RM L52G25

A PRELIMINARY INVESTIGATION OF SHOCK-WAVE
REFLECTIONS IN A SMALL CLOSED BALLISTIC
RANGE WITH VARIOUS TYPES OF WALLS. A. P-
Sabol. September 1952. 21p. photos., diagrs.
(NACA RM L52G25) (Declassified from Confidential,
10/12/54)







NACA
RESEARCH ABSTRACTS NO.74


all important respects. The addition of chamber and
twist reduced the maximum lift-drag ratios over
mos~t of the speed range tested but increased the
maximum lift coefficient at all speed*



NACA RM L52H11

A STUDY OF THE FLOW FIELD BEHIND THE TRI.
ANGULAR HORIZONTAL TAIL OF A CANARD AIR-
PLANE AT APPROXIMATELY THE VERTICAL-
TAIL LOCATION BY MEANS OF A TUFT GRID.
Joseph L. Johnson, Jr. October 1952. 18p. diagrs.,
tab. (NACA RM L52H11) (Declassified from
Confidential, 10/12/54)

An investigation of the flow field behind the horizon-
tal triangular tail of a canard model by means of a
tuft grid placed at approximately the vertical-tail
location (about 6.0 horizontal-tail root chords behind
the horizontal tail) indicated that trailing vortices
from the horizontal tail produced a sidewash field
over the model which probably accounted for the vari-
ation in the static directional stability and damping-
in-yaw characteristics of canard models reported in
previous investigations,



NACA RM L52H20

A TRANSONIC INVESTIGATION OF THE AERO-
DYNAMIC CHARACTERISTICS OF PLATE- AND
BELL-TYPE OUTLETS FOR AUXILIARY AIR.
William J. Nelson and Paul E. Dewey. September
1952. 25ip. diagrs., photos. (NACA RM L52H20)
(Declassified from Confidential, 10/12/54)

The aerodynamic characteristics of several plate-
type and bell-mouthed outlets representative of cur-
rent design practice have been investigated at Mach
nubers from 0.7 to 1.3. These data show that the
effect of an external stream on the discharge coef-
ficient of such outlets is determined primarily by
the rate of discharge relative to the free-jet flow
rate. For outlets of the types tested, low-speed
data may probably be applied at transonic Mach num-
bers if the free-jet characteristics of the outlet are
first corrected for the effects of Reynolds number
and Mach number on the discharge characteristics
of the outlet discharging into infinite space. Total-
pressure surveys through the jet wake are compared
at Mach numbers of 0.7, 1. 1, and 1.3 for the dif-
ferent outlets investigated. Static-pressure distri-
bution in the vicinity of circular outlets is shown for
several Mach numbers and discharge rates.



NACA RM L52H21

INVESTIGATION OF THE, VARIATION WITH
REYNOLDS NUMBER OF THE BASE, WAVE, AND
SKIN-FRICTION DRAG OF A PARABOLIC BODY OF
REVOLUTION (NACA RM-10) AT MACH NUMBERS
OF 1.62, 1.93, AND 2.411N THE LANGLEY 9-INCH
SUPERSONIC TUNNEL. Eugene S. Love, Donald E.
Co~letti and August F. Bromm, Jr. October 1952.
62p. diagrs., photos., (NACA RM L52H21)
(Declassified from Confidential, 10/12/54)


Results are presented from an investigation at
M = 1.62, 1.93, and 2.41 of the variation with
Reynolds number of the base, wave, and skin-
friction drag of a parabolic body of revolution (NACA
RM-10). Comparisons are made with theory and
other experimental data. An empirical expression
relating the Reynolds n-mber of transition to Mach
number is presented as well as an explanation of the
behavior of base pressure with varying Reynolds
number.



NACA RM L52IO2

INVESTIGATION AT HIGH AND LOW SUBSONIC
MACH NUMBERS OF TWO SYMMETRICAL 6-
PERCENT-THICK AIRFOIL SECTIONS DESIGNED
TO HAVE HIGH MAXIMUM LIFT COEFFICIENTS
AT LOW SPEEDS. Nicholas J. Paradiso. October
1952. 37p. diagrs., photo., 2 tabs. (NACA
RM LS2102) (Declassified from Confidential,
10/12/54)

Results are presented of an investigation at both
high and low subsonic Mach numbers of two newly
derived symmetrical 6-percent-thick airfoil sections
which are of a family of airfoil sections designed for
high maximum lift at low subsonic speeds. As com-
pared to previously tested sections of this family,
the sections of this investigation have less blunt
leading edges. Included are results for Reynolds
numbers up to 9 x 106 and for two locations of sur-
face roughness.



NACA RM L52I03

THRUST LOADING OF THE. NACA 3-(3)(05)-05
EIGHT-BLADE DUAL-ROTATING PROPELLER AS
DETERMINED FROM WAKE SURVEYS. Robert J.
Platt, Jr. October 1952. 44rp. diagrs., photo.
(NACA RM L52103) (Declassified from Confidential,
10/12/54)

Wake-survey measurements of the thrust loading of
the NACA 3-(3)(05)-05 eight-blade dual-rotating
propeller are presented to supplement the previously
published force-test results for this propeller. The
data cover a blade-angle range from 650 to 800,
measured at 0.75R, at forward Mach numbers from
0.35 to 0.925.



NACA RM LS2IO8

FORCE TESTS OF THREE THIN WINGS OF MOD-
ERATELY LOW ASPECT RATIO AT HIGH SUB-
SONIC MACH NUMBERS. Gareth H. Jordan.
October 1952. 22p. diagrs. (NACA RM L52108)
(Declassified from Confidential, 10/12/54)

Results are presented of force tests made in the
Langley 24-inch high-speed tunnel on three thin
wings of moderately low aspect ratio at high subsonic
Mach numbers to determine the effect of leading-
edge shape and section profile on the aerodynamic
characteristics. The range of angle of attack was
from -2o to 8o and the range of Mach number was
from 0.30 to about 0.90.







NACA
RESEARCH ABSTRACTS NO. 74

An unswe~pt-wing airplane and a 350 swept-wing air-
plane were flown in rough air to investigate effects
of sweep on gust loads and gust selectivity. The
swept-wing airplane experienced lower loads in
rough air than the unswept-wing airplane. The ratio
of loads on the two airplanes was equal to the cosine
of the angle of sweep and to the ratio of the lift-curve
slopes from either low-speed wrind-tunnel data or
calculated from the empirical formula 6A cos A
A + 2 cos2n
where A is aspect ratio and A is the angle of
swe.Only small differences were shown in the
gust selectivity characteristics of the two airplanes.


NACA RM L52LO9

WING AND FUSELAGE LOADS MEASURED IN
FLIGHT ON THE NORTH AMERICAN B-45 AND
F-82 AIRPLANES. Paul W. Harper. February
1953. 35p. diagrs., 4 tabs. (NACA RM LS2LO9)
(Declassified from Confidential, 10/12/54)

Flight investigations were conducted to determine
the wing and fuselage loads on the B-45 and F-82
airplanes by means of calibrated strain-gage installa-
tions at each wing- and tail-fuselage juncture. The
tests covered a Mach number range of approximately
0.3 to 0.75. The aerodynaic loads measured for
the B-45 airplane were substantially as predicted by
theory, but the loads on the F-82 airplane were in
disagreement with theory. A gradual outboard shift
in wing center of pressure with increasing Mach
number was noted for both airplanes.



NACA RM L52L11

EXPERIMENrTAL INVESTIGATION OF THE FLOW
FIELD BEHIND AN ASPECT-RATIO-10 HYDROFOIL
NEAR THIE WATER SURFACE. Arthur WI. Carter
and Roger V. Butler. February 1953. 31p. diagrs.,
photos., tab. (NACA RM L52L11) (Declassified from
Confidential, 10/12/54)

An investigation was made at subcritical speeds of
the flow field behind an aspect-ratio-10 hydrofoil
operating at a depth below the free-water surface of
0.75 chord. The downwash and water surface pro-
files were measured behind the hydrofoil over a
range of lateral and longitudinal positions of interest
for tandem hydrofoil applications. The experimental
data were compared with theoretical predictions
based on two-dimensional flow.


NACA RM L52L22

FREE-FLIGHT-TUNNEL INVESTIGATION OF THE
LOW-SPEED STABILITY AND CONTROL CHIAR-
ACTERISTICS OF A MODEL HAVING A FUSELAGE:
OR RELATIVELY FLAT CROSS SECTION. John W.
Paulson and Joseph L. Johnson, Jr. February 1953.
30p. diagrs., photo., tab. (NACA RM L52L22)
(Declassified from Confidential, 10) 12/54)

Results are presented of an experimental investiga-
tion in the Langley free-flight tunel to determine
the dynamic lateral stability and control character-
istics of a model having a relatively flat cross-
section fuselage. Tests were made with several
vertical-tail configurations and with the leading-edge
flap retracted and extended.


NACA RM L52IO9

ANALYTICAL STUDY OF STATIC AND LOW-
SPEED PERFORMANCE OF THIN PROPELLERS
USING TWO-SPEED GEAR RATIOS TO OBTAIN
OPTIMUM ROTATIONAL SPEEDS. Jean Gilman, Jr.
November 1952. 52p. diagrs., 4 tabs. (NACA
RM L52109) (Declassified from Confidential,
10/12/54)

This paper presents methods of estimating the static
and low-speed performance of thin propellers when
operating at optimum rotational speeds. The im-
portance, under certain conditions, of incorporating
variable gearing is illustrated by specific examples.
The effect of camber on static thrust is also inves-
tigated, and it is shown that thin propellers having
moderately cambered blade sections can produce 15
to 20 percent more static thrust than propellers
having symmetrical blade sections.



NACA RM L52K18a

A STUDY OF THE: USE OF VARIOUS HIGH-LIFT DE-
VICES ON THE HORIZONTAL TAIL OF A CANARD
AIRPLANE MODEL AS A MEANS OF INCREASING
"THE ALIDWABLE CENTER-OF-GRAVITY
TRAVEL. Joseph L.Johnson, Jr. January 1953.
25p. diagrs., 2 tabs. (NACA RM L52K18a)
(Declassified from Confidential, 10/12/54)

This paper contains results of a low-speed power-
off static longitudinal stability and control investiga-
tion to study the use of various high-lift devices on
the horizontal tail of a canard airplane model as a
means of increasing the allowable center-of-gravity
travel.



NACA RM L52K28a

AVERAGE SKIN-FRICTION COEFFICIENTS FROM
BOUNDARY-LAYER MEASUREMENTS ON A OGIVE-
CYLINDER BODY IN FLIGHT AT SUPERSONIC
SPEEDS. J. Dan Loposer. January 1953. 11p.
diagrs., photo. (NACA RM LS2K28a) (Declassified
from Confidential, 10/12/54)

Boundary-layer measurements on a rocket-powered
free-flight model to determine average skin-friction
coefficients have been made on an ogive-cylinder
body of fineness ratio 15.9. Average skin-friction
coefficients were obtained for the body area ahead of
the fins over a range of Mach number from 1.3 to
2.5 and over a range of Reynolds number from
90.3 x 106 to 162.9 x 106 (based on axial body length
to the measurement station). Comparison of the
experimental data with a flat-plate skin-friction
theory by Van Driest showed good agreement.



NACA RM L52LO2

A COMPARISON OF GUST LOADS MEASURED IN
FLIGHT ON A SWEPT-WING AIRPLANE AND AN
UNSWEPT-WING AIRPLANE. Jack Funk and
Harry C. Mickleboro. June 1953. 16p. diagrs.,
2 tabs. (NACA RM L52LO2) (Declassified from
Confidential, 10/12/54)







NACA
RESEARCH ABSTRACTS NO. 74

NACA RM L52L26a

EPFFCTS OF ROUGHNESS AND REYNOLDS NUM-
BER ON THE NONLINEAR LIFT CHARACTERISTICS
OEF WING WfI~t MODIcFIED HbEXAGONAL AIRFOIL


rM L2169a5)3(D~e~ca a~itd frpoh Co Aie~nial,

Nonlinear lift characteristics of a low-aspect-ratio
wing with modified hexagonal airfoil sections at low
Reynolds numbers were observed at Mach numbers
from 0.370 to 0.896. Effects of increased Reynolds
numbers and of surface roughness on the linearity of
the lift curves were investigated.


NACA RM L53A19

SOME TORSIONAL-DAMPING MEASUREMENTS OF
LAMINATED BEAMS AS APPLIED TO THE
PROPELLER STALL-FLUTTER PROBLEM.
Atwood R. Heath, Jr. April 1953. 14p. diagrs.,
photos., tab. (NACA RM L53A19) (Declassified from
Confidential, 10/12/54)

The structural damping in the torsion mode of vibra-
tion of a series of untwisted, laminated thin beams
simulating propeller blades is presented. The num-
ber of laminations were varied, as well as the bond-
ing material and the method of joining laminations.
Application of the data to the calculation of the mini-
mum flutter speed of thin propeller blades indicates
that appreciable gains in the minimum flutter speed
may be obtained for laminated blades using a
Cycleweld bond,


NACA RM L53B19

WIND-TUNNEL INVESTIGATION OF THE EFFECTS
OF VARIOUS DORSAL-FIN AND VERTICAL-TAIL
CONFIGURATIONS ON THE DIRECTIONAL STABIL-
ITY OF A STREAMLINED BODY OF TRANSONIC
SPEEDS. TRANSONIC-BUMP METHOD. Harold S.
Johnson and William C. Hayes. April 1953. 22p.
diagrs., photo., tab. (NACA RM L53B19) (Declas-
sified from Confidential, 10/12/54)

Yawing-moment coefficients were obtained for
several dorsal-fin and vertical-tail configurations in
combination with a streamlined body through a large
angle-of-sideslip range and a Mach number range of
0.59 to 1.11. The results indicated that dorsal fins
improved the directional stability characteristics of
the body alone and the body--vertical-tail configura-
tion. A ring tail was more effective at small angles
of sidealip and less effective at large angles of side-
slip than a tapered low-aspect-ratio vertical tail.


NACA RM L53C10

A PRELIMI~INARY INVESTIGATION OF AERODY.
NAMIC CHARACTERISTICS OF SMALL INCLINED
AIR OUTLETS AT TRANSONIC MACH NUMBERS.
Paul E. Dewey. April 1953. 21p. photos., diagrs.
(NACA RMl L53C10) (Declassified from Confidential,
10/12/54)

The aerodynamic characteristics of several outlets


with inclined or curved axes discharging air into a
transonic stream. have been investigated. The data
presented herein show the discharge coefficient of
such outlets and static-pressure distribution in the
vcinitnuof t outlets forrseveral values of stream

observations, showing the vortex formation caused
bythe outlet discharge from a perpendicular and an



NACA RM L53D21

MEASUREMENTS OF AERODYNAMLIC CHAR-
ACTERISTICS AT TRANSONIC SPEEDS OF AN
UNSWEPT AND UNTAPERED NACA 65-009 AIR-
FOIL MODEL OF ASPECT RATIO 3 WITH 1/4-
CHORD PLAIN FLAP BY THE NACA WING-FLOW
METHOD. Harold I. Johnson. June 1953. 35p.
diagrs., photo. (NACA RM L53D)21) (Declassified
from Confidential, (10/12 54)

Lift, pitching-moment, and hinge-moment data ob-
tained from wing-flow tests of an unswept, untapered
NACA 65-000 airfoil model of aspect ratio 3.01
equipped with 1/4l-chord full-span plain flap are pre-
sented. Effects of flap gap and of roughness were
Investigated. The Mach number range of the tests
was 0.65 to 1.10 and the Reynolds number range was
0.5 x 106 to 0.9 x 106

NACA RM L53D30

CALCULATION OF AERODYNAMIC FORCES ON AN
INCLINED DUAL-ROTATING PROPELLER. John
L. Crigler and Jean Gilman, Jr. June 1953. 24p.
diagrs. (NACA RM L53D30) (Declassified from
Confidential, 10/12/54)

This paper presents a method of calculating the
fluctuating aerodynamic forces on a dual-rotating
propeller the thrust axis of which is inclined at an
angle to the air stream. Sample calculations are
made and the results are analyzed to show some of
the effects encountered with this type of operation.


NACA RM L53E12

EFFECTS OF COMPRESSIBILITY AT MACH NUM-
BERS UP TO) 0.8 ON INTERNAL-FLOW CHARAC-
TERISTICS OF A COWLING-SPINNER COMBINA-
TION EQUIPPED WIH AN EIGHT-BLADE DUAL-
ROTATION PROPELLER. Gene J. Bingham and
Arvid L. Keith, Jr. June 1953. 39p. diagrs.,
photos. (NACA RM L53E12) (Declassified from
Confidential, 10/12/54)

An investigation has been conducted to study the
effects of compressibility for Mach nubers; up to
0.8 on the internal-flow characteristics of an NACA
1-series cowling-spinner combination equipped with
a dual-rotation propeller. Two propellers having
24-percent-thick shank sections were studied; one
had the propeller shanks extended to the spinner sur-
tace and the juncture sealed, and the other had a
raised platform-type juncture with the gap required
to allow blade-angle changes located outside of the
spinner boundary layer. The effects of variations in
inlet height and rate of internal compression on the
internal-flow characteristics were also studied.
Total- and static-pressure distributions and average
impact pressure coefficients measured at the inlet
and at a diffuser station are presented.







NACA
RESEARCH ABSTRACTS NO.74

NACA RM L53G23

TRANSONIC AERODYNAMIC CHARACTERISTICS
OF AN NACA 64A006 AIRFOIL SECTION WITH A
15-PERCENT-CHORD LEADING-EDGE FLAP.
Milton D. Humphreys. September 1953. 44p.
diagrs., photos. (NACA RM L53G23) (Declassified
from Confidential, 10/12/54)

Two-dimensional airfoil section normal-force, drag,
pitching-moment, flap normal-force, and hinge-
moment characteristics obtained on an NACA
64A006 airfoil section equipped with a 15-percent-
chord leading-edge flap are presented for Mach num-
bers from 0.5 to 1.0. The leading-edge flap ef-
fected a general improvement in the force charac-
teristics at Mach numbers up to 0.8 by alleviating
flow separation at high normal-force coefficients.



NACA RM L53G23a

THE EFFECT OF CONTROL-SURFACE-SERVO
NATURAL FREQUENCY ON THE DYNAMIC PER-
FORMANCE CHARACTERISTICS OF AN ACCELER-
ATION CONTROL SYSTEM APPLIED TO A SUPER-
SONIC MISSILE. Anthony L. Passera and Martin L.
Nason. September 1953. 28p. diagrs., 3tabs.
(NACA RM L53G23a) (Declassified from
Confidential, 10/12/54)

A theoretical investigation is made to determine the
effect of the natural frequency of a second-order
control-surface servo upon the dynamic performance
characteristics of a normal-acceleration control
system with respect to accuracy of control and accu-
mulator volume of flow and peak rate of volume flow
for several missile scale sizes and flight conditions
in response to a step-input command. This paper
seeks a compromise value of natural frequency that
yields a control system with moderate volume flow
and peak rate of volume flow along with good
accuracy of control.






NACA RM L53H13

THE AERODYNAMIC CHARACTERISTICS AT
TRANSONIC SPEEDS OF AN ALL-MOVABLE,
TAPERED, 450 SWEPTBACK, ASPECT-RATIO-4
TAIL DEFLECTED ABOUT A SKEWED HINGE AXIS
AND EQUIPPED WITHAN INSET UNBALANCING
TAB. James M. Watson. September 1953. 40p.
diagrs. (NACA RM L53H13) (Declassified from
Confidential, 10/12/54)

This paper contains the results of an investigation at
transonic speeds of an all-movable, aspect-ratio-4,
taper-ratio-0.6 tail swept back 450 at the quarter-
chord line, and deflected about a skewed hinge axis
located behind the centers of pressure. The tail was
equipped with an inset tab. Lift, pitching-moment,
and hinge-moment data are presented for various tail
angles of attack and deflections and tab deflections
through a Mach number range of 0.61 to 1.21 obtained
by the transonic-bump technique.


NACA RM L53E28a

INVESTIGATION AT TRANSONIC SPEEDS OF THE
HINGE-MOMENT AND LIFT-EFFECTIVENESS
CHARACTERISTICS OF A SINGLE FLAP AND A
TANDEM FLAP ON A 600 DELTA WING. Delwin
R. Croom and Harleth G. Wiley. July 1953. 16p.
diagrs. (NACA RM L53E28a) (Declassified from
Confidential, 10/12/54)

This paper presents the results of an investigation
to determine the comparative hinge-moment and lift-
effectiveness characteristics of a single flap and a
tandem flap on a semispan 600 delta-wing model at
transonic speeds by the transonic-bump method.
The delta-wing model was a flat plate with beveled
leading and trailing edges, a maximum thickness
ratio of 0.045, 600 sweepback at the leading edge, a
taper ratio of 0, and an aspect ratio of 2.31. Lift
and hinge moments were obtained through a Mach
number range of 0.60 to 1.11. The tandem flap had
less variation of Ch6 (hinge-moment coefficient per
degree flap deflection) with Mach number than did
the single flap and the lift effectiveness was only
about 50 percent of that obtained with the single flap.




NACA RM L53Fl6a

LOW-SPEED INVESTIGATION OF THE LATERAL
CONTROL CHARACTERISTICS OF THREE TIP AIL-
ERONS ON A 600 TRIANGULAR WING. Stanley M.
Gottlieb. August 1953. 24p. diagrs., photo. (NACA
RM L53F16a) (Declassified from Confidential
10/12/54) '

An investigation has been made at a Reynolds num-
ber of 9 x 106 and a Mach number of 0.15 of the
lateral control characteristics of three wing-tip
ailerons on a 6-percent-thick, 600 triangular-wing-
fuselage combination. The controls consisted of
two half-delta ailerons having areas equal to 0.077
and 0.138 times the wing-semispan area and a full-
delta aileron having an area equal to 0.138 times the
wing-semispan area. The tests, which included
measurements of rolling moments, hinge moments,
yawing moments, and lateral force, were made
through ranges of angle of attack and deflection
from -120 to 380 and from -20" to 200, respectively.


NACA RM L53GO8

INVESTIGATION TO DETERMINE EFFECTS OF
RECTANGULAR VORTEX GENERATORS ON THE
STATIC-PRESSURE DROP THROUGH A 900 CIRCU-
LAR ELBOW. E. Floyd Valentine and Martin R.
Copp. September 1953. 35p. diagrs., photos.
(NACA RM L53G08) (Declassified from Confiden-
tial, 10/12/54)

An investigation was made of a constant-area, circu-
lar 900 elbow of mean radius or curvature equal to
its diameter with several arrangements of simple,
nontwisted, rectangular vortex generators. They
were installed at the inlet and also at stations 15o
300, and 600 into the elbow. The effect of the vortex
generators on the pressure drop measured between
the inlet and a station 4 diameters downstream of
the elbow is given for one inlet-boundary-layer
thickness.


NACA-Lanley 11-30-54 4M5












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