Research abstracts

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

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Subjects / Keywords:
Aeronautics -- Abstracts -- Periodicals   ( lcsh )
Aeronautics -- Research -- Abstracts -- Periodicals   ( lcsh )
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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:
AA00009235:00010

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National Advisory Committee for Aeronautics


Research Abstracts


NO.50


OCTOBER 9, 1953


CURRENT NACA REPORTS

NACA Rept. 1073

AN ITERATIVE TRANSFORMATION PROCEDURE
FOR NUMERICAL SOLUTION OF FLUTTER AND
SIMILAR CHARACTERISTIC-VALUE PROBLEMS.
Myron L. Gossard. 1952. ni, 45p. diagrs.. 9
tabs. (NACA Rept. 1073. Formerly TN 2346)

The idea of the iterative transformation procedure
suggested by H. Wielandt is explained. Application
of the procedure to ordinary natural-vibral ion
problems and to flutter problems is shown in numer-
ical examples. Comparisons of computed results
with experimental values and with results obtained
by other methods of analysis are, made.

NACA Rept. 1077 .. -

TWO- AND THREE-DIMENSIONAI UNSTS BD9 f
LIFT PROBLEMS IN HIGH-SPEE F
Harvard Lomax, Max A. Heaslet, ranklyn B.
Fuller and Loma Sluder. 1952. ii, 55p. diagrs.,
3 tabs. (NACA Rept. 1077. Formerly TN 2403;
TN 2387)

The problem of transient lift on two- and three-
dimensional wings flying at high speeds is discussed
as a boundary-value problem for the classical wave
equation. Kirchhoff's formula Is applied so that the
analysis is reduced, just as in the steady state, to
an investigation of sources and doubles. The appli-
cations include the evaluation of indicial lift and
pitching-moment curves for two-dimensional sinking
and pitching wings flying at Mach numbers equal to
0, 0.8, 1.0, 1.2, and 2 0. Results for the sinking
case are also given for a Mach number of 0.5. In
addition, the indicial functions for supersonic-
edged triangular wings in both forward and reversed
flow are presented and compared with the two-
dimensional values.


NACA Rept. 1080

A THEORETICAL ANALYSIS OF THE EFFECTS OF
FUEL MOTION ON AIRPLANE DYNAMICS. AJbertl' t&-
A. Schy. 1952. ii, 22p. diagrs., 2 tabs. (NakGA-
Rept. 1080. Formerly TN 2280)

The general equations of motion for an airplane with
a number of spherical fuel tanks are derived. These
equations are applied to two cases with two fuel tanks
located in the plane of symmetry. The calculated
motions show that the airplane motion may be
greatly changed by considering the motion of the fuel
and, in particular, that small-amplitude residual
oscillations may result. The same type of analysis
may be applied to arbitrarily shaped tanks; there-

*AVAILABLE ON LOAN ONLY.


fore, the most general conclusions as to the effects
of the fuel motion on airplane dynamics also apply
for arbitrarily shaped tanks.

NACA Rept. 1090

METHOD FOR CALCULATING LIFT DISTRIBU-
TIONS FOR UNSWEPT WINGS WITH FLAPS OR
AILERONS BY USE OF NONLINEAR SECTION LIFT
DATA. James C. Sivells and Gertrude C.
Westrick. 1952. ii, 25p. diagrs. 13 tabs.
(NACA Rept. 1090. Formerly TN 2283)

A method is presented for calculating lift distribu-
tions for unswept wings with flaps or ailerons
uaing nonlinear section lift data. This method is
based upon lifting-line theory and is an extension to
thS method described in NACA.Rep'85-... Simpli-
J fie& computing forms cont4A4ngdlaed.fulinples
are ven for both symme(rjirand asymetfacal
S J distributions. A fe roknparisons of expIri-
mental and calculated ah cteristics ar aso.
presented. k l3 -J 1


NACA Rept. 1101 /

FLIGHT INVESTIGATIONOF tE CA-L
FEEL DEVICE IN AN IRREVERSIBLE ELEATOR
CONTROL SYSTEM OF A LARGE-MfRPtANE.
B. Porter Brown, Robert G. Chilton and James B.
Whitten. 1952. ii, 14p. diagrs. (NACA
Rept. 1101. Formerly TN 2496)

Data are presented showing the flight characteristics
of a large airplane having a control-surface booster
and mechanical feel device in the elevator-control
system. The tests were made with various force
gradients provided by the adjustable feel device.
;- The booster was set to operate at a very high boost
S, ratio throughout the tests so that the measured or
9 j3pparjnt stick-free stability would be influenced only
slightly by the aerodynamic hinge moments. The
Sr sultslshow the effect of the feel device on the
ha dling qualities of the test airplane and also the
design features which should be incorporated in such
Mel devices.

NACA Rept. 1105

CHORDWISE AND COMPRESSIBILITY CORREC-
TIONS TO SLENDER-WING THEORY. Harvard
Lomax and Loma Sluder. 1952. ii, 19p. diagrs.,
4 tabs. (NACA Rept. 1105. Formerly TN 2295)

Corrections to slender-wing theory are obtained by
assuming a spanwise distribution of loading and de-
termining the chordwise variation which satisfies the
appropriate integral equation. Such integral equa-
tions are set up in terms of the given vertical in-
duced velocity on the center line or, depending on the


ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST, NW, WASHINGTON s25. D. C., CITING CODE NUMBER ABOVE EACH TITLE,
THE REPORT TITLE AND AUTHOR.
I Of 2-
*U.5--e







2

type of wing plan form, its average value across the
span at a given chord station. The chordwise dis-
tribution is then obtained by solving these integral
equations. Results are shown for flat-plate,
rectangular, and triangular wings.


NACA Rept. 1107

AN EMPIRICALLY DERIVED BASIS FOR CALCU-
LATING THE AREA, RATE, AND DISTRIBUTION
OF WATER-DROP IMPINGEMENT ON AIRFOILS.
Norman R. Bergrun. 1952. ii, 21p. diagrs.,
6 tabs. (NACA Rept. 1107)

An empirically derived basis for predicting the area,
rate, and distribution of water-drop impingement on
airfoils of arbitrary section is presented. The
concepts involved represent an initial step toward
the development of a calculation technique which is
generally applicable in the design of thermal ice-
prevent ion equipment for airplane wing and tail
surfaces. The calculation technique presented is
based on results of extensive water-drop trajectory
computations for five airfoil cases which consisted
of 15-percent-thick airfoils encompassing a moder-
ate lift-coefficient range. The differential equations
pertaining to the paths of the drops were solved by
a differential analyzer.


NACA Rept. 1109

EXPERIMENTAL INVESTIGATION OF BASE PRES-
SURE ON BLUNT-TRAILING-EDGE WINGS AT
SUPERSONIC VELOCITIES. Dean R. Chapman,
William R. Wimbrow and Robert H. Kester. 1952.
ii, 19p. diagrs., photos., tab. (NACA Rept. 1109.
Formerly TN 2611)

The pressures acting on the base of blunt-trailing-
edge airfoils have been measured at Mach numbers
of 1.25, 1.5, 2.0, and 3.1 and at Reynolds numbers
from 0.2 to 3.8 million. Data are presented for 29
profiles both with laminar and with turbulent bound-
ary layers approaching the trailing edges of the
wings. The base pressure is found to be a function
primarily of Mach number and the ratio of the bound-
ary layer thickness at the trailing edge to the
trailing-edge thickness.

NACA Rept. 1111

AN ANALYSIS OF LAMINAR FREE-CONVECTION
FLOW AND HEAT TRANSFER ABOUT A FLAT
PLATE PARALLEL TO THE DIRECTION OF THE
GENERATING BODY FORCE. Simon Ostrach.
APPENDIX B: NUMERICAL SOLUTION OF SIM-
PLIFIED BOUNDARY-VALUE PROBLEM. Lynn U.
Albers. 1953. ii, 17p. diagrs., tab. (NACA
Rept. 1111. Formerly TN 2635)

A formal and general analysis of the free-convection
flow about a flat plate parallel to the direction of the
generating body force is made, and velocity and tem-
perature distributions for Prandtl numbers of 0.01,
0.72, 0.733, 1, 2, 10, 100, 1000, and large Grashof
numbers are computed. The distributions for
Prandtl number of 0.72 compare favorably with ex-
perimental values. It is shown that velocities and
Nusselt numbers of the same order of magnitude as
those associated with forced-convection flows can be


NACA
RESEARCH ABSTRACTS NO.50

obtained under free-convection conditions. A flow
and a heat-transfer parameter are derived from
which the important physical quantities can be com-
puted. Reasonable agreement is obtained among
values of the heat-transfer parameter obtained from
an approximate theoretical development, experi-
ments, and the present development.

NACA TN 2979

EFFECTS OF SYMMETRIC AND ASYMMETRIC
THRUST REVERSAL ON THE AERODYNAMIC
CHARACTERISTICS OF A MODEL OF A TWIN-
ENGINE AIRPLANE. Kenneth W. Goodson and
John W. Draper. September 1953. 67p. diagrs.,
photo., tab. (NACA TN 2979)

An investigation was made to determine the magni-
tude and degree of changes in static forces and
moments caused by variations of symmetric and
asymmetric thrust reversal on a twin-engine air-
plane. The effects of both positive and negative
thrust coefficients were investigated on the Indi-
vidual propellers and covered a thrust-coefficient
range of 0.167 to -0.150.


NACA TN 2983

LINEARIZED POTENTIAL THEORY OF PRO-
PELLER INDUCTION IN A COMPRESSIBLE FLOW.
Robert E. Davidson. September 1953. 47p. diagrs.,
5 tabs. (NACA TN 2983)

This paper gives the potential-theory representation
of the wave-equation flow about a lifting-line pro-
peller of finite number of blades and arbitrary cir-
culation distribution. From the velocity potential,
the compressible inflow velocities at the blade be-
came known. The induced velocities are known also
at any point in the flow because the velocity potential
is determined for the whole field.


NACA TN 2997

APPLICATION OF SEVERAL METHODS FOR DE-
TERMINING TRANSFER FUNCTIONS AND FRE-
QUENCY RESPONSE OF AIRCRAFT FROM FLIGHT
DATA. John M. Eggleston and Charles W.
Mathews. September 1953. 74p. diagrs., 2 tabs.
(NACA TN 2997)

A study is presented of several methods for deter-
mining the transfer functions and frequency response
of aircraft from flight tests. Results obtained from
experience in the use of these methods are com-
pared as to time required for application, compara-
tive accuracy, and means for facilitating their use.
The studies cover three categories of methods:
sinusoidal response, Fourier analysis of transients,
and curve-fitting analysis of transients Three
general types of aircraft are used to illustrate the
application of these methods.


NACA TN 2999

IMPINGEMENT OF DROPLETS IN 900 ELBOWS
WITH POTENTIAL FLOW. Paul T. Hacker.
Rinaldo J. Brun and Bemrose Boyd. September
1953. 58p. diagrs., 2 tabs. (NACA TN 29991






NACA
RESEARCH ABSTRACTS NO.50


Trajectories were determined for droplets in air
finiii, t&hroupi 900 elbows especially designed for
two-dimensional potential motion with low pressure
losses. The elbows were established by selecting
as walls of each elbow two streamlines of the flow
field produced by a complex potential function that
establishes a two-dimensional flow around a 900
bend. An unlimited number of elbows with slightly
differ r ti shapes can be established by selecting
different pairs of streamlines as walls. The elbows
produced by the complex potential function selected
are suitable for use in aircraft air-intake ducts.
The droplet impingement data derived from the
trajectories are presented along with equations in
such a manner that the collection efficiency, the
area, the rate, and the distribution of droplet im-
pngernement can be determined for any elbow de-
fined by any pair of streamlines within a portion of
the flow field established by the complex potential
function. Coordinates for some typical stream-
lines of the flow field and elocirt, components for
several points along these streamlines are pre-
sented in tabular form.


NACA TN 3005

HEAT TRANSFER AND SKIN FRICTION BY AN
INTEGRAL METHOD IN THE COMPRESSIBLE
LAMINAR BOUNDARY LAYER WITH A STREAM-
WISE PRESSURE GRADIENT. Ivan E. Beckwith.
September 1953. 55p. diagrs., tab. (NACA
TN 30051i

A simplified method has been developed for the cal-
culation of heat transfer and skin friction in the
compressible laminar boundary layer with an arbi-
trary Prandtl number near unity and an arbitrary
streamwise pressure gradient and wall temperature
distribution. The use of a fifth-degree polynomial
for the stagnation enthalpy profile gives results of
good accuracy for the case of boundary-layer
cooling. The method has been extended to the calcu-
lation of heat transfer under conditions of equili-
brium dissociation in the boundary layer. By means
of a suitable transformation the method is also
applied to a body of revolution with a boundary-
layer thickness of the order of the body radius.


NACA TN 3007

LIFT AND PITCHING MOMENT AT LOW SPEEDS
OF THE NACA 64A010 AIRFOIL SECTION
EQUIPPED WITH VARIOUS COMBINATIONS OF A
LEADING-EDGE SLAT, LEADING-EDGE FLAP,
SPLIT FLAP, AND DOUBLE-SLOTTED FLAP.
John A. Kelly and Nora-Lee F. Hayter. September
1953. 45p. diagrs., photos., 2 tabs. (NACA
TN 3007)

Results of a two-dimensional wind-tunnel investiga-
tion of an NACA 64A010 airfoil equipped with a
leading-edge flap, a leading-edge slat, a split flap,
and a double-slotted flap are presented. The re-
sults include determination of opt lnun-m slat posi-
tions and effects of varying Reynolds number on the
lift and pitching-moment characteristics of the
model with the various high-lift devices.


3


NACA TN 3008

EFFECTS OF FINITE SPAN ON THE SECTION
CHARACTERISTICS OF TWO 450 SWEPTBACK
WINGS OF ASPECT RATIO 6. Lynn W. Hunton.
September 1953. 32p. diagrs. (NACA TN 3008.
Formerly RM A52A10)

A study has been made of the Finite -s.pri effects on
the local loading characteristics of two siepitrack
wings at low speed with a view toward providing some
insight into the usefulness of two-dimensional section
data and span-loading theory for determining the sec-
tion characteristics of a swept wing. The two wings
considered were identical in plan form having 450 of
sweepback of the quarter-chord line, an aspect ratio
of 6, and a taper ratio of 0.5 but differed in twist and
in sections, the latter being the NACA 64A010 and
NACA 64A810. The analysis is based on compari-
sons of local pressure distributions and local lift
characteristics on the wings with comparable two-
dimensional section data, all of which were available
at large scale.

NACA TN 3009

VELOCITY POTENTIAL AND AIR FORCES
ASSOCIATED WITH A TRIANGULAR WING IN
SUPERSONIC FLOW, WITH SUBSONIC LEADING
EDGES, AND DEFORMING HARMONICALLY
ACCORDING TO A GENERAL QUADRATIC EQUA-
TION. Charles E. Watkins and Julian H. Berman.
September 1953. 61p. diagrs., tab. (NACA
TN 3009)

The velocity potential for a triangular wing with sub-
sonic leading edges experiencing harmonic deforma-
tions in supersonic flow is treated herein. The
oscillations considered are such that the amplitude of
distortion of the wing can be represented by a general
quadratic equation. The velocity potential is derived
in the form of a power series in terms of the fre-
quency of oscillation. Although only the first four
terms of the series expansion are presented, addi-
tional terms may be obtained if desired. The mate-
rial constitutes an extension of the work given in
NACA Report 1099.

NACA TN 3011

COEFFICIENT OF FRICTION AND DAMAGE TO
CONTACT AREA DURING THE EARLY STAGES OF
FRETTING. I GLASS, COPPER, OR STEEL
AGAINST COPPER. Douglas Godfrey and John M.
Bailey. September 1953. 23p. diagrs., photos.,
2 tabs. (NACA TN 3011)

Experiments were conducted to measure the coeffi-
cient of friction 1i and to determine the damage to
the contact area during early stages of fretting of
copper at a frequency of 5 cycles per minute.
Specimen combinations of copper against glass,
copper against copper, and copper against steel, as
well as various copper oxide films and powder com-
pacts, were used. The results lead to the conclu-
sion that fretting of copper starts with the same
mechanical damage that occurs during unidirectional
sliding. Fretting of copper against glass, copper
against copper, and copper against steel starts with
adhesion and metal transfer (galling) with accompa-
nying high g values (>1.0) the same as those ob-
tained during unidirectional sliding. After the ini-







4

tial high values of ji, a reduction in 1 was ob-
served, associated with reduced plowing and an in-
creasing concentration of debris in and around the
contact area. After approximately 100 cycles of
fretting, gi reached a constant value (0.5-0.6) ap-
proximately the same as that obtained with com-
pacts of either cuprous or cupric oxide. The
presence of preformed cuprous or cupric oiide films
on copper does not delay the occurrence of fretting
but only lowers the initial coefficient of friction.

NACA TN 3012

AN ANALYSIS OF TURBOJET-ENGINE-INLET
MATCHING. DeMarquis D. Wyatt. September
1953. 19p. diagrs. (NACA TN 3012)

A method of presenting turbojet-engine air-flow
requirements and inlet-system air-flow capacities
in identical though independent parametric terms is
developed. The application of the air-flow repre-
sentation technique to the analysis of engine-inlet
matching conditions is demonstrated. Several ex-
amples are presented to illustrate the application of
the method to the explicit determination of inlet
geometric variations required to improve the power-
plant performance of supersonic airplanes.

NACA TN 3014

CALCULATED SPANWISE LIFT DISTRIBUTIONS
AND AERODYNAMIC INFLUENCE COEFFICIENTS
FOR UNSWEPT WINGS IN SUBSONIC FLOW.
Franklin W. Diederich and Martin Ziotnick.
September 1953. 120p. diagrs., 11 tabs. (Tables
of F matrices to be used with TN 3014 are published
separately) (NACA TN 3014)

Spanwise lift distributions have been calculated for
nineteen unswept wings with various aspect ratios
and taper ratios and with a variety of angle-of-
attack or twist distributions, including flap and aile-
ron deflections, by means of the Weissinger method
with eight control points on the semispan. Also
calculated were aerodynamic influence coefficients
which pertain to a certain definite set of stations a-
long the span, and several methods are presented for
calculating aerodynamic influence coefficients for
stations other than those stipulated.


NACA TN 3015

AN EXPERIMENTAL INVESTIGATION OF SECOND-
ARY FLOW IN AN ACCELERATING, RECTANGULAR
ELBOW WITH 900 OF TURNING. John D. Stanitz,
Walter M. Osborn and John Mizisin. October 1953.
60p. diagrs., photos., 2 tabs. (NACA TN 3015)

Secondary flow tests were conducted on an accelerat-
ing elbow with 900 of turning designed for prescribed
velocities that eliminate boundary-layer separation
by avoiding local decelerations along the walls. Sec-
ondary flows were investigated for six boundary-layer
thicknesses generated on the plane walls of the elbow
by spoilers upstream of the elbow inlet. The passage
vortex associated with secondary flows appears to be
near the suction surface and away from the plane
wall of the elbow at the exit and does not have appre-
ciable spanwise motion as it moves downstream from
the elbow exit. As the spoiler size increases, the
boundary-layer form changes and a rather sudden


NACA
RESEARCH ABSTRACTS NO.0S -


difference in the secondary flow occurs, perhaps
associated with the reduced importance of viscous
effects in thick boundary layers. It is suggested that
the strength of the secondary vortices is small and
that the energy of secondary flows is small.

NACA TN 3016

ANALYSIS OF TURBULENT HEAT TRANSFER AND
FLOW IN THE ENTRANCE REGIONS OF SMOOTH
PASSAGES. Robert G. Deissler. October 1953.
88p. diagrs. (NACA TN 3016)

A previous analysis for fully developed turbulent heat
transfer and flow with variable fluid properties is
extended and applied to the entrance regions of
smooth tubes and parallel flat plates. Integral hbeat-
transfer and momentum equations are used for calcu-
lating the thicknesses of the thermal and flow bound-
ary layers. The effect of variable properties is
determined for the case of uniform heat flux, uniform
initial temperature distribution, and fully developed
velocity distribution. A number of other cases In
which the fluid properties are,constant are analyzed.
The predicted Nusselt numbers for air with a uniform
wall temperature and uniform initial temperature and
velocity distributions agree closely with experimen-
tally determined values.


NACA TN 3017

AXIAL-LOAD FATIGUE TESTS ON NOTCHED AND
UNNOTCHED SHEET SPECIMENS OF 61S-T6
ALUMINUM ALLOY, ANNEALED 347 STAINLESS
STEEL, AND HEAT-TREATED 403 STAINLESS
STEEL. Herbert F. Hardrath, Charles B. Landers
and Elmer C. Utley, Jr. October 1953. 28p.
diagrs., 4 tabs. (NACA TN 3017)

Axial-load fatigue tests at a stress ratio of zero
were performed on notched and unnotched sheet
specimens of 61S-T6 aluminum alloy and 347 and
403 stainless steels. Special emphasis was placed
on tests at high stress levels producing failures in
small numbers of cycles. It was found that the
stress-concentration factors effective in fatigue of
notched specimens were somewhat less than the
theoretical elastic values at low stresses and were
approximately equal to one at the ultimate strength.
The mimimum life to failure at stresses near the
ultimate strength was drastically reduced with m-
creasing stress-concentration factor.


NACA TN 3019

INVESTIGATION OF THE STATISTICAL NATURE
OF THE FATIGUE OF METALS. G. E. Dieter and
R. F. Mehl. Carnegie Institute of Technology.
September 1953. 25p. diagrs., 5 tabs. (NACA
TN 3019)

Results are presented of an investigation of the sta-
tistical nature of the fatigue of metals utilizing sta-
tistical methods developed previously. The inves-
tigation included a study of the fatigue properties and
their statistical variation of a plain carbon eutectoid
steel heat-treated to coarse pearlitic and spheroi-
dized structures of the same tensile strength and of
commercially pure aluminum (2S) and 245 alloy
heat-treated to two different structures. Calcula-







NACA
RESEARCH ABSTRACTS NO.50


tion of data from the literature provided statistics
for 75S aluminum alloy for comparison with the
data of the present investigation.


NACA RM E53E05

EFFECT OF PRESSURE ON THE SMOKING TEND-
ENCY OF DIFFUSION FLAMES. Rose L. Schalla
and Glen E. McDonald. September 1953. 13p.
diagrs., 2 tabs. (NACA RM E53E05)

The effect of pressure on smoke formation was in-
vestigated by burning nine hydrocarbon fuels as
diffusion flames from a modified wick lamp in an en-
closed chamber. The maximum relative rate at
which each fuel could be burned without smoking was
determined over a pressure range of about 1/2 to 4
atmospheres, and up to 12 atmospheres in the case
of one fuel. The results indicate that over this
pressure range the maximum smoke-free fuel flow
was inversely proportional to the pressure. The
relative variation in smoking tendency among the
different fuel types was approximately constant at
all pressures. From an analysis of the data it has
been tentatively proposed that the variations in
smoke formation with pressure result from changes
in the rate of diffusion and mixing of the fuel and
air.


NACA RM E53G08

VAPOR PRESSURES OF CONCENTRATED NITRIC
ACID SOLUTIONS IN THE COMPOSITION RANGE
83 TO 97 PERCENT NITRIC ACID, 0 TO 6 PER-
CENT NITROGEN DIOXIDE, 0 TO 15 PERCENT
WATER, AND IN THE TEMPERATURE RANGE 200
TO 800 C. A. B. McKeown and Frank E. Belies.
September 1953. 22p. diagrs., tab. (NACA
RM E53G08)

Total vapor pressures were measured for 28 acid
mixtures of the ternary system nitric acid, nitrogen
dioxide, and water within the temperature range 200
to 800 C and within the composition range 83 to 97
percent nitric acid, 0 to 6 percent nitrogen dioxide,
and 0 to 15 percent water. The ullage of the appara-
tus used for the measurements was 0. 65. Ternary
diagrams showing isobars as a function of composi-
tion of the system NO2-H20-HNO3 were constructed
from experimental and interpolated data for the
temperatures 250, 400, and 600 C and are presented
herein.

NACA RM L53G24a

MODEL DITCHING INVESTIGATIONS OF THREE
AIRPLANES EQUIPPED WITH HYDRO-SKIS.
(Revised) Lloyd J. Fisher. September 1953. 8p.
photos. (NACA RM L53G24a)

Calm-water tests were made to determine possible
arrangements of hydro-ski ditching gear on typical
multiengine airplanes. The tests showed that hydro-
skis would afford very satisfactory water landings
as compared with landings without skis. The best
Landings were made in a near-level (slightly nose-
up) landing attitude although any normal landing
attitude was satisfactory. It is possible that critical
damage could be eliminated from ditchings by using
a hydro-ski landing gear.


5


BRITISH REPORTS


N-11456*

Aeronautical Research Council (Gt. Brit.)
THE DISTURBED MOTION OF ARTICULATED
BLADES. H. Roberts. October 1949. 77p.
diagrs. (ARC 12, 688; H. 127)

The general theory of motion for articulated heli-
copter blades is presented and allowance is made for
for: (a) arbitrary motion of the hub, (b) offset blade
hinges, (c) inclined flapping hinges, and (d) cam-
bered blades. The method employed is the simple
one of giving the hub prescribed horizontal, vertical,
and angular velocities, the motion being two-
dimensional. It is shown that the flapping stability
equation arises when considering disturbed motion
or response phenomena. It is also shown that dis-
turbances in the pitching plane lead in general to a
lateral tilt of the rotor disk, thereby inducing a
coupling between the longitudinal and lateral
stability equations.


N-26520

Aeronautical Research Council (Gt. Brit.)
THE DETERMINATION OF SKIN TEMPERATURES
ATTAINED IN HIGH SPEED FLIGHT. F. V.
Davies and R. J. Monaghan. 1953. 65p. diagrs.,
8 tabs. (ARC CP 123)

This report discussed the factors affecting skin
temperatures attained by high-speed missiles and
presents some methods of solution. These have
been reduced to graphical or tabular form and are
set out in order of complexity. Graphical or alge-
braic solutions may be quickly obtained if steady con-
ditions are assumed, and for some flight cases
these are reasonable approximations to correspond-
ing transient solutions. If the temperature time
variation is required, then longer numerical integra-
tion processes have to be performed. Account may
be taken of external radiation and heat loss to the in-
terior if their effects are considered significant.


N-26521*

Aeronautical Research Council (Gt. Brit.)
THE DEFINITIONS OF THE ANGLES OF INCI-
DENCE AND OF SIDESLIP. C. H. E. Warren.
1953. lip. diagrs. (ARC CP 124)

The use of large angles of incidence and of sideslip
in missile work, and recent changes in wind-tunnel
testing techniques, .have shown the need for clear
and precise definitions of the angles of incidence and
of sideslip. The suitability of different definitions
for both experimental and theoretical work in both
the aircraft and missile field is considered, and it
is concluded that, as no single definition is univer-
sally acceptable, care should be taken in theoretical
and experimental reports to define precisely the
angles used.







6


N-26522"

Aeronautical Research Council (Gt. Brit.)
PRESSURE ERROR MEASUREMENT USING THE
FORMATION METHOD. K. C. Levon. 1953. 15p.
diagrs. (ARC CP 126)

Measurements of pressure error at altitude have
been made by flying several aircraft in formation
with a reference aircraft whose airspeed system had
previously been calibrated by radar. Tests made
show that analysis by comparison of indicated air-
speeds (comparison of differences between static and
total head pressures) gives more consistent and
reliable results than pressure altitude comparison
(comparison of direct measurement of static pres-
sure).


N -26523*

Aeronautical Research. Council (Gt. Brit.)
SURFACE SLOPES AND CURVATURES OF THE
RAE 100 104 AND OTHER ROOFTOP-TYPE
AEROFOIL SECTIONS. J. Williams and Edna M.
Love. 1952. lip. 3 tabs. (ARC CP 129)

Formulas and tables have been obtained for the
accurate and rapid calculation of the surface slopes
and curvatures of the RAE 100-104 airfoil sections,
and of more general "rooftop-type" sections. In
particular, the surface slopes of the RAE 102 and
104 shapes have been evaluated for application with a
"tangent-plane" method of model construction.

N-26535*

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF COMPRESSIBILITY ON THE
PERFORMANCE OF A GRIFFITH AEROFOIL.
H. H. Pearcey and E. W. E. Rogers. 1953. 30p.
diagrs., photos., 4 tabs. (ARC R & M 2511.
Formerly ARC 10, 096; FM 1017; Perf. 250)

Experiments were made in the 20-inch by 8-inch
high-speed tunnel at the NPL on a 9-inch chord, 22
percent thick, symmetrical Griffith section at 00
incidence. Drag was determined by the pitot
traverse method. Information on the flow was ob-
tained from the pitot tube traverses, from direct
shadow photographs, and from normal pressure
measurements. Three Mach numbers of the un-
disturbed stream were covered, namely 0.4, 0.6, and
0.7. Estimates of the power absorbed by the com-
pressor, ignoring duct losses are made from meas-
urements of the mass of air sucked and the static
pressure in the slots. Additional information was
obtained on the adverse effect of a large radius of the
forward lip of the slot, on the effect at Mo = 0.4 of
an increase in slot width, and on the choke quantities
for the slots.


N-26536 *

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF SLIPSTREAM ON THE LONGI-
TUDINAL STABILITY OF MULTI-ENGINED AIR-
CRAFT. D. E. Morris and J. C. Morrall. 1953.
9p. diagrs., tab. lARC R& M 2701; ARC 12.136.
Formerly RAE Aero 2304)


NACA
RESEARCH ABSTRACTS NO.50


Flight measurements of longitudinal stability power-
off and power-on made on numerous aircraft have
been analyzed and a generalized curve for estimating
the contribution of slipstream to longitudinal
stability. applicable to both flaps-up and flaps-down
cases, has been derived. Using this curve, the
change in stability due to slipstream at a given value
of CL can be estimated with a probable error of
less than t 0.02 in the position of the neutral point.

N-26537 *

Aeronautical Research Council (Gt. Brit.)
LOW-SPEED WIND-TUNNEL TESTS ON TWO 45
DEG SWEPTBACK WINGS OF ASPECT RATIOS 4.5
AND 3.0 (MODELS A AND B). J. Trouncer and G. F.
Moss. 1953. 43p. diagrs., 13 tabs.
(ARC R & M 2710; ARC 10, 904. Formerly
RAE Aero 2210)

Low-speed stability tests were made on two wings of
aspect ratio 4.5 and 3.0. Both wings were of 450
sweepback, 4:1 taper ratio, and 14 percent thickness
ratio. Tests included stability tests on the two wings
without body or tail unit, tests with a body, fin and
tailplane fitted and tests made with two types of nose
flaps on the aspect ratio 4.5 wing.


N-26538*

Aeronautical Research Council (Gt. Brit.)
THE MEASUREMENT OF THE OVERALL DRAG OF
AN AIRCRAFT AT HIGH MACH NUMBERS. D. J.
Higton, R. H. Plascott and D. A. Clarke. 1953.
22p. diagrs., photo., tab. (ARC R & M 2748;
ARC 11,429; ARC 12,237. Formerly RAE Aero
2241; RAE Aero 2309)

This report describes the technique which has been
developed to measure the over-all drag of an air-
craft at high Mach numbers in both level flight and
dives. It shows how improvements have been made
both in flight and tunnel technique so that compari-
sons between full-scale and model tests have now
become possible. Flight results from Meteor IV
aircraft show close agreement between drag
measured in level flight and in dives and later tests
compare well with high-speed wind-tunnel meas-
urements on a 1/12th scale model.


N-26540*

Aeronautical Research Council (Gt. Brit.)
THE THEORETICAL PRESSURE DISTRIBUTIONS
AROUND SOME CONVENTIONAL TURBINE BLADES
IN CASCADE. T. J. Hargesi 1953. lOp. diagrs.
(ARC R & M 2765; ARC 13,360. Formerly NGTE
R. 67)

By means of Relf's analogy between aerodynamic
streamline flow and electric potential flow, the
theoretical pressure distributions around a series of
conventional turbine blades in cascade have been
determined over a range of incidence covered in
some previously reported aerodynamic tests. The
theoretical pressure distributions and their variation
with incidence provide the basis of an explanation of
the observed aerodynamic performance.






NACA
RESEARCH ABSTRACTS NO.50 7

N-26541* transition and separation. A brief consideration is
also given to some of the effects of airstream turbu-
Aeronautical Research Council (Gt. Brit.) lence, weak shock waves from the walls, and heat
A CALCULATION OF THE COMPLETE DOWNWASH transfer.
IN THREE DIMENSIONS DUE TO A RECTANGULAR
VORTEX. Doris E. Lehrian. 1953. 45p. diagrs.,
15 labs. (ARC R&M 2771. Formerly ARC 11,786;
S L. C 2252; Perf. 487; ARC 12, 221; S & C 2294;
Perf. 537)

A calculation of the complete downwash in three di-
mensions due to a rectangular vortex, is given for
the limited range Z = +4. The downwash is com-
puted at selected positions, in planes normal to the
plane of the vortex; these planes are spaced at even
integral multiples of the semiwidth of the vortex,
measured from the line of symmetry. Values are
tabulated for Z in the range (0, 4) and a set of
graphs is also included for 0 < Z < 2; they are to be
used in conjunction with the "Tables of Complete
Downwash due to a Rectangular Vortex"
(R. & M. 2461).


MISCELLANEOUS


NACA TN 2991

Errata No. 1 on "ACCELERATIONS AND PASSEN-
GER HARNESS LOADS MEASURED IN FULL-
SCALE LIGHT-AIRPLANE CRASHES." A. Martin
Elband, Scott H. Simpkinson and Dugald 0. Black.
August 1953.



UNPUBLISHED PAPERS


N-20019A*


REVIEW OF PUBLISHED DATA ON THE EFFECT
OF ROUGHNESS ON TRANSITION FROM LAMINAR
TO TURBULENT FLOW. Hugh L. Dryden. 6p.
diagrs. (Reprint from Journal of the Aeronautical
Sciences, v. 20, no. 7, July 1953, p. 477-482).

A review is presented of the published data on the
effect of roughness, especially single roughness
elements, on transition from laminar to turbulent
flow, in which an attempt is made to reanalyze and
correlate the available information. The paper also
discusses available data on the effect of distributed
roughness on transition on a flat plate, as well as
some of the published data on roughness effects on
transition on airfoils.


N-25045*


SOME FACTORS CONTRIBUTING TO SCALE
EFFECT AT SUPERSONIC SPEEDS. Ira H. Abbott.
34p. diagrs., photos. (Presented at the fourth
meeting of the Advisory Group for Aeronautical
Research and Development, Wind Tunnel and Model
Testing Panel, London, September 3-11, 1953)

This paper presents information on the Reynolds
number effects at supersonic speeds on skin friction,


NACA-Lanldev 10-9-53 4M





UNIVERSITY OF FLORIDA


3 1262 08153 274 8




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