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
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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:
AA00009235:00086

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Full Text

National Advisory Committee for Aeronautics


Research Abstracts


NO.25


JULY 1, 1952


CURRENT NACA REPORTS


NACA Rept. 1014
STUDY OF EFFECTS OF SWEEP ON THE FLUTTER
OF CANTILEVER WINGS. I. G. Barmby, H. J.
unuitngham and L. E. Garrick. 1951. ii, 25p. diagrs.,
:p ihoto., 7 tabs. (NACA Rept. 1014. Formerly
TN 2121; RM L8H30)

S As experimental and analytical investigation of the
flutter of uniform, cantilever, sweptback wings is
Reported. The experiments employed group of wings
Jept back by rotating and by shearing. The angle of
.*i"eop ranged from 00 to 60 and the Mach numbers
tond ed to approximately 0.85. Comparison with
experiment indicates that the analysis developed in
L the present paper is satisfactory, at least for nearly
:i Wform cantilever wings of moderate length-chord


N... A Rept. 1018

A1:aEORETICAL ANALYSIS OF THE EFFECT OF
....3 ii f LAG IN AN AUTOMATIC STABILIZATION
C filf ::pTXM ON THE LATERAL OSCILLATORY STA-
NUT?^::: 01BaTOF AN AIRPLANE. Leonard Sternfleld and
dwailp y B. Gates, Jr. 1951. ii, 12p. diagrs. (NACA
(||t|:i'.iW 1018. Formerly TN 2005)
., ..... A method is presented for determining the effect of
..t:ae lag... in an automatic stabilization system on the
/ flatra oscillatory stability of an airplane. The
Pw: thp is applied to a typical present-day airplane
etuIped with an automatic pilot sensitive to yawing
aftleration and geared to the rudder so that rudder
ceAtrol is applied in proportion to the yawing accel-
eAtion.


NACA Rept. 1027

BUCKLING OF THIN-WALLED CYLINDER UNDER
AXIAL COMPRESSION AND INTERNAL PRESSURE.
Hsu Lo, Harold Crate and Edward B. Schwartz. 1951.
Op. diagrs. (NACA Rept. 1027. Formerly TN 2021)

An investigation was made of a thin-walled cylinder
under axial compression and various internal pres-
sures to study the effect of the internal pressure on
the compressive buckling stress of the cylinder. A
theoretical analysis based on a large-deflection theory
was also made. The theoretically predicted increase
of compressive buckling stress due to internal pres-
sure agrees fairly well with the experimental results.


NACA Rept. 1031

A STUDY OF THE USE OF EXPERIMENTAL STA-
BILITY DERIVATIVES IN THE CALCULATION OF
THE LATERAL DISTURBED MOTIONS OF A SWEPT-
WING AIRPLANE AND COMPARISON WITH FLIGHT
RESULTS. John D. Bird and Byron M. Jaquet. 1951.
ii, 25p. diagrs., photos., 4 tabs. (NACA Rept. 1031.
Formerly TN 2013)

Experimentally determined lateral-stability deriva-
tives of a swept-wing airplane are presented to show
effects of slots, flaps, propeller, and ventral fins and
are used in calculations of a number of lateral dis-
turbed motions. The calculated motions are com-
pared with those obtained in flight tests to determine
the applicability of the experimental stability deriva-
tives. The effects of nonlinearity of some of the
aerodynamic forces with sideslip are shown by a few
supplemental calculations.


NACA Rept. 1036

EXPERIMENTAL INVESTIGATION OF THE
EFFECTS OF VISCOSITY ON THE DRAG AND
BASE PRESSURE OF BODIES OF REVOLUTION AT
A MACH NUMBER OF 1.5. Dean R. Chapman and
Edward W. Perkins. 1951. ii, 24p. photos., diagrs.
(NACA Rept. 1036. Formerly NACA RM A7A31a)

Models were tested to evaluate effects of Reynolds
number for both laminar and turbulent boundary
layers. Principal geometric variables investigated
were afterbody shape and length-diameter ratio.
Force tests and base-pressure measurements were
made. Schlieren photographs were used to analyze
the effects of viscosity on flow separation and shock-
wave configuration and to verify the condition of the
boundary layer as deduced from the force tests. The
results are discussed and compared with theoretical
calculations.


NACA Rept. 1039

ON THE PARTICULAR INTEGRALS OF THE
PRANDTL-BUSEMANN ITERATION EQUATIONS
FOR THE FLOW OF A COMPRESSIBLE FLUID.
Carl Kaplan. 1951. ii, 6p. (NACA Rept. 1039.
Formerly TN 2159)

The particular integrals of the second-order and
third-order Prandtl-Busemann iteration equations
for the flow of a compressible fluid are obtained by
means of the method in which the complex conjugate
variables z and z are utilized as the independent
variables of the analysis. The assumption is made
that the Prandtl-Glauert solution of the linearized or
first-order iteration equation for the two-dimensional


AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASH I NGTON 2s, D. C.. CITING CODE NUMJBER, TlITLE AND AUTHOR.
Ei '::' :.


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2


flow of a compressible fluid is known. The forms of
the particular integrals, derived for subsonic flow,
are readily adapted to supersonic flows with only a
change in sign of one of the parameters of the prob-
lem.


NACA TN 2489

AERODYNAMIC CHARACTERISTICS OF A REFINED
DEEP-STEP PLANING-TAIL FLYING-BOAT HULL
WITH VARIOUS FOREBODY AND AFTERBODY
SHAPES. John M. Riebe and Rodger L. Naeseth.
June 1952. 48p. photos., diagrs., 8 tabs. (NACA
TN 2489. Formerly RM L8F01)

An investigation was made in the Langley 300-mph
7- by 10-foot tunnel to determine the aerodynamic
characteristics of a refined deep-step planing-tail
hull with various forebody and afterbody shapes and,
for comparison, a streamline body simulating the
fuselage of a modern transport airplane. The results
of the tests indicated that the configurations incor-
porating a forebody with a length-beam ratio of 7 had
lower minimum drag coefficients than the configura-
tions incorporating a forebody with length-beam ratio
of 5. The lowest minimum drag coefficients, which
were considerably less than that of a conventional
hull and slightly less than that of a streamline body,
were obtained on the length-beam-ratio-7 forebody,
alone and with round center boom. Drag coefficients
and longitudinal- and lateral-stability parameters
presented include the interference of a 21-percent-
thick support wing.



NACA TN 2641

A VECTOR STUDY OF LINEARIZED SUPERSONIC
FLOW APPLICATIONS TO NONPLANAR PROBLEMS.
John C. Martin. June 1952. 81p. diagrs., tab. (NACA
TN 2641)

A vector study of the partial-differential equation of
steady linearized supersonic flow is presented. Gen-
eral expressions are derived which relate the veloc-
ity potential in the stream to the conditions on the
disturbing surfaces. Problems concerning nonplanar
systems are investigated, and methods are derived
for the solution of some simple problems. The damp-
ing in roll is found for rolling tails consisting of four,
six, and eight rectangular fins.



NACA TN 2657

SOME EFFECTS OF FREQUENCY ON THE CON-
TRIBUTION OF A VERTICAL TAIL TO THE FREE
AERODYNAMIC DAMPING OF A MODEL OSCIL-
LATING IN YAW. John D. Bird, Lewis R. Fisher
and Sadie M. Hubbard. April 1952. 39p. diagrs.,
photo., tab. (NACA TN 2657)

The damping in yaw and the directional stability of
a model freely oscillating in yaw were measured tail-
off and tail-on and compared with the values obtained
by theoretical consideration of the unsteady lift
associated with an oscillating vertical tail. A range


NACA
RESEARCH ABSTRACTS NO.25


of low frequencies comparable to those of the lateral
motions of airplanes was covered. The analysis in-
cludes the effects of vertical-tail aspect ratio and
the two-dimensional effects of compressibility.


NACA TN 2670

HIGH-SPEED SUBSONIC CHARACTERISTICS OF 16
NACA 6-SERIES AIRFOIL SECTIONS. Milton D.
Van Dyke. March 1952. 65p. diagrs., tab. (NACA
TN 2670. Formerly RM A7J23)

High-speed subsonic characteristics have been meas-
ured for NACA 63-, 64-, 65-, and 66-series airfoil
sections having thickness ratios of 6, 8, 10, and
12 percent and an ideal lift coefficient of 0. 2. For
given thickness ratio, airfoil sections with minimum
pressure near 40-percent chord exhibit the best over-
all aerodynamic characteristics. For even the
thinnest sections, good high-speed performance is
maintained over a wide range of lift coefficient.


NACA TN 2672

THEORETICAL AUGMENTATION OF TURBINE-
PROPELLER ENGINE BY COMPRESSOR-INLET
WATER INJECTION, TAIL-PIPE BURNING, AND
THEIR COMBINATION. Reece V. Hensley. March
1952. 43p. diagrs. (NACA TN 2672)

A theoretical evaluation of the performance of tur-
bine propeller engines with augmentation by water
injection, tail-pipe burning, and by their combina-
tion was made. The investigation covered the flight
conditions for which each method of augmentation was
most suitable for Mach numbers up to 1.1 and alti-
tudes up to 35,000 feet. The variation in augmenta-
tion with varying compressor and turbine efficiencies,
compressor pressure ratio, and turbine-inlet temper-
ature was calculated. The augmentation with varying
ambient temperature and relative humidity was deter-
mined and the dependence of the augmentation on the
completeness of evaporation during compression was
investigated.


NACA TN 2676

SUMMARY OF STALL-WARNING DEVICES. John A.
Zalovcik. May 1952. 15p. diagrs. (NACA TN 2676)

Principles involved in the operation of stall-warning
devices are described and conditions under which
difficulty may be experienced are pointed out. Some
specific examples of stall-warning devices are illus-
trated and described.


NACA TN 2689

EFFECT OF HIGH-LIFT DEVICES ON THE LOW-
SPEED STATIC LATERAL AND YAWING STABILITY
CHARACTERISTICS OF AN UNTAPERED 45 SWEPT-
BACK WING. Jacob H. Lichtenstein. May 1952.
20p. diagrs., photo. (NACA TN 2689. Formerly
RM L8G20)







NACA
RESEARCH ABSTRACTS NO.25


Results of a low-speed wind-tunnel investigation to
determine the effect of high-lift devices on the static
lateral stability derivatives and the yawing derivatives
of an untapered 45 sweptback wing are presented.
The tests were made in the curved-flow test section
of the Langley stability tunnel at a Reynolds number
of 1.1 x 106.


NACA TN 2692

ON THE FORM OF THE TURBULENT SKIN-
FRICTION LAW AND ITS EXTENSION TO COM-
PRESSIBLE FLOWS. Coleman duP Donaldson. May
1952. 19p. diagrs. (NACA TN 2692)

A derivation of the form of the incompressible turbu-
lent skin-friction law for an insulated flat plate is
made in such a way that it may be extended to com-
pressible flows. The ratio of compressible to incom-
pressible skin friction is obtained, and the results
are shown to be in good agreement with existing ex-
perimental results.


NACA TN 2695

MIGRATION OF COBALT DURING FIRING OF
GROUND-COAT ENAMELS ON IRON. William N.
Harrison, Joseph C. Richmond, Joseph W. Pitts and
Stanley G. Benner, National Bureau of Standards.
June 1952. 28p. photos., 5 tabs. (NACA TN 2695)

A typical porcelain-enamel ground coat containing a
small amount of radioactive cobalt as oxide was pre-
pared, applied to sandblasted enameling-iron blanks,
and fired under conditions selected to give underfired,
normally fired, and overtired coatings. Examination
of these specimens revealed that a cobalt-bearing
metallic deposit had been formed at the enamel-metal
interface durijig firing, and that the enamel layer was
depleted in cobalt oxide near the interface. The
amounts of such deposit and depletion were found to
vary directly with the severity of the firing treatment.
The amount of metallic deposit was extremely small,
on the order of 0.01 micron thick if computed as a
continuous layer of metallic cobalt in the case of the
normally fired specimens.


NACA TN 2697

METHOD AND GRAPHS FOR THE EVALUATION OF
AIR-INDUCTION SYSTEMS. George B. Brajnikoff.
April 1952. 41p. diagrs., tab. (NACA TN 2697)

Graphs that allow rapid evaluation of air-induction
systems from considerations of their aerodynamic
parameters in combination with power-plant charac-
teristics are presented for the supersonic Mach
numbers up to 3.0. Restrictions imposed by the
engine characteristics on the use of a fixed-size air
inlet are discussed and illustrated by means of
sample solutions. The relation between the engine
characteristics, flight conditions, inlet character-
istics, and inlet area for optimum performance is
given.


NACA TN 2698

THEORETICAL ANALYSIS OF HYDRODYNAMIC
IMPACT OF A PRISMATIC FLOAT HAVING FREE-
DOM IN TRIM. Robert W. Miller. June 1952. 31p.
diagrs., 2 tabs. (NACA TN 2698)

The paper presents the derivation and a method for
solution of equations taking into account the effects of
freedom in trim on the loads, moments, and motions
of a prismatic, V-bottom float forebody during hydro-
dynamic impact. Comparisons of solutions of the
equations with fixed-trim solutions and with some ex-
perimental data are presented and indicate that the
solutions are reasonable and consistent with both the
experimental data and with general experience.


NACA TN 2699

CALCULATION OF LIFT AND PITCHING MOMENTS
DUE TO ANGLE OF ATTACK AND STEADY PITCH-
ING VELOCITY AT SUPERSONIC SPEEDS FOR THIN
SWEPTBACK TAPERED WINGS WITH STREAMWISE
TIPS AND SUPERSONIC LEADING AND TRAILING
EDGES. John C. Martin, Kenneth Margolis and
Isabella Jeffreys. June 1952. 116p. diagrs., 9 tabs.
(NACA TN 2699)

On the basis of linearized supersonic-flow theory
formulas for the stability derivatives Cm CL and
Cmq were derived for a series of thin sweptback
tapered wings with streamwise tips and supersonic
leading and trailing edges. The results of the analy-
sis are presented as a series of design charts. Some
illustrative variations of the derivatives and of the
chordwise center-of-pressure location with the vari-
ous wing design parameters are also included.


NACA TN 2702

AN APPROXIMATE METHOD OF DETERMINING
THE SUBSONIC FLOW IN AN ARBITRARY STREAM
FILAMENT OF REVOLUTION CUT BY ARBITRARY
TURBOMACHINE BLADES. Chung-Hua Wu, Curtis
A. Brown and Vasily D. Prian. June 1952. 46p.
diagrs., 4 tabs. (NACA TN 2702)

A method is presented to obtain a relatively quick
approximate solution of the detailed subsonic flow of
a nonviscous fluid past arbitrary turbomachine blade
sections lying on an arbitrary surface of revolution.
Calculation Is first made for the flow along a particu-
lar streamline about midway between two adjacent
blades. Extension of the flow from this streamline to
the blade surfaces is then made by the use of a
Taylor's series, and a method is provided to correct
the solution to fit the given blade shape and operating
conditions. The method is illustrated with examples
of compressible flow in a turbine cascade and a
centrifugal compressor. Three terms of the
Taylor's series are found to give sufficient accuracy.
Sufficient convergence is obtained for the turbine cas-
cade after two cycles of computation and for the
centrifugal compressor after four cycles of compu-









4


station. (Each cycle of computation takes 16 hours.)
The detailed flow distribution obtained compares
very well with an available numerical solution and
experimental data.


NACA TN 2703

ELECTRICAL TECHNIQUES FOR COMPENSATION
OF THERMAL TIME LAG OF THERMOCOUPLES
AND RESISTANCE THERMOMETER ELEMENTS.
Charles E. Shepard and Isidore Warshawsky. May
1952. ii, 85p. diagrs., photos., 9 tabs. (NACA TN
2703)

Basic electrical networks are described that com-
pensate for the thermal time lag of thermocouple and
resistance thermometer elements. For a given set
of operating conditions, networks requiring no ampli-
fiers can provide a thirtyfold reduction in effective
time lag. This improvement is obtained without
attenuation of the voltage signal, but does result in a
large reduction in the amount of electric power
available because of an increase in the output imped-
ance of the network. Networks using commercially
available amplifiers can provide a thousandfold re-
duction in the effective time lag without attenuation of
the alternating voltage signal or of the available
electric power, but the improvement is often obtained
at the expense of loss of the zero-frequency signal.
The completeness of compensation is limited by the
extent of off-design operation required.


NACA TN 2704

FATIGUE STRENGTHS OF 14S-T4 ALUMINUM
ALLOY SUBJECTED TO BIAXIAL TENSILE
STRESSES. Joseph Marin and W. P. Hughes,
Pennsylvania State College. June 1952. 24p.
photos., diagrs., 5 tabs. (NACA TN 2704)

The purpose of this investigation was to determine
the influence of biaxial tensile stresses on the fatigue
strength of a 14S-T4 aluminum alloy when subjected
to various ratios of biaxial stresses. The effect
upon the fatigue strength of varying the ratios of bi-
axial stresses was studied. The biaxial fatigue
stresses were produced by applying simultaneously a
pulsating internal pressure and a fluctuating axial
tensile load to a thin-walled tubular specimen. The
maximum and minimum values of the longitudinal and
circumferential stresses were kept in phase by adjust-
ing the testing machine. The fatigue strengths and
S-N diagrams were obtained up to about 107 cycles
for four principal stress ratios.



NACA TN 2705

THEORY OF SUPERSONIC POTENTIAL FLOW IN
TURBOMACHINES. Robert H. Wasserman. June
1952. 44p. diagrs. (NACA TN 2705)

A general method for solving supersonic potential
flow problems for stationary or rotating coordinate
systems is presented. The principal attributes of
the method are: It can handle flows which cannot be


NACA
RESEARCH ABSTRACTS NO.25


treated as two-dimensional, and a sound theoretical
basis gives assurance of its validity for a class of
boundary-value problems. An application to the
design of a compressor rotor is made.


NACA TN 2706

EFFECT OF CHANGING PASSAGE CONFIGURATION
ON INTERNAL-FLOW CHARACTERISTICS OF A
48-INCH CENTRIFUGAL COMPRESSOR. I CHANGE
IN BLADE SHAPE. Donald J. Michel, John Mizisin
and Vasily D. Prian. May 1952. 40p. diagrs., photo.,
tab. (NACA TN 2706)

The passage contour of a 48-inch centrifugal impeller
was modified by changing the shape of the blades. A
comparison of the internal-flow characteristics was
made of the original and modified passage contours at
the design flow rate. In addition, the internal-flow
characteristics of the modified passage are presented
and analyzed over the entire flow range from maximum
flow to surge at a corrected impeller tip speed of
700 feet per second. At design flow, the modified
impeller, which had lesser deceleration rates along
the blade surfaces, showed a general improvement in
efficiency throughout the passage over that of the
original impeller. At high weight flows (negative
angles of attack), there were large separation losses
at the inlet because of the shape of the leading edge;
at low flow rates (positive angle of attack), the sepa-
ration losses were not as great. The low-efficiency
regions in the passage generally occurred near the
flow surfaces with decelerating flow, except along the
driving face where most boundary-layer build-up by
adverse velocity gradients was probably eliminated by
secondary flows or bleeding through the clearance
space.


NACA TN 2707

ANALOGUE-COMPUTER SIMULATION OF AN AUTO-
PILOT SERVO SYSTEM HAVING NONLINEAR RE-
SPONSE CHARACTERISTICS. Arthur L. Jones and
John S. White. June 1952. 30p. diagrs. (NACA
TN 2707)

The nonlinear response characteristics of an electro-
hydraulic servo system were successfully simulated
using an electronic analogue computer (differential
analyzer). In obtaining a satisfactory simulation it
was necessary not only to take into account an easily
recognized nonlinear amplifier response but also the
accumulative effect of some time lags of the servo
system.


NACA TN 2708

COMPARISON OF THREE MULTICYLINDER ICING
METERS AND CRITIQUE OF MULTICYLINDER
METHOD. Wallace E. Howell, Mount Washington
Observatory. June 1952. 40p. diagrs., photos.,
6 tabs. (NACA TN 2708)

Three multicylinder icing meters, fundamentally
similar but differing from each other in important








NACA
RESEARCH ABSTRACTS NO.25


deisgn details, were compared in use at the Mount
Washington Observatory. Comparison of relative
effectiveness of the instruments, evaluation of obser-
vational errors, determination of the effects of de-
S tailed design differences, and recommendations for
further improvements of design are presented. An
j evaluation of the multicylinder method, concerned
With the validity of the theoretical basis and the de-
gree to which the instruments and the technique of
their use permit accurate determinations of the phys-
ical measurements involved, is also included.


NACA TN 2709

FATIGUE AND STATIC TESTS OF FLUSH-RIVETED
JOINTS. Darnley M. Howard and Frank C. Smith,
National Bureau of Standards. June 1952. 38p.
photos., diagrs., 3 tabs. (NACA TN 2709)

Fatigue and static tests were made of 1 8-inch diam-
eter: A17S-T3 100 countersunk-head rivets in lap and
butt Joints. Both machine-countersunk and dimpled
boles were used. The sheet materials were 0.032-
inch-thick bare and alclad 24S-T3 and 75S-T6 alumi-
num alloys and 0.064-inch-thick alclad 75S-T6 alloy.
.From the results no satisfactory single relation be-
tween static properties and fatigue life covering the
four materials could be found.


NACA TN 2710

DIFFUSION OF HEAT FROM A LINE SOURCE IN
ISOTROPIC TURBULENCE. Mahinder S. Uberoi and
Stanley Corrsin, Johns Hopkins University. June
1952. 90p. dlagrs., photos., tab. (NACA TN 2710)

An experimental and analytical study has been made
of some features of the turbulent heat diffusion be-
hind a line heated wire stretched perpendicular to a
flowing isotropic turbulence. The mean temperature
distributions have been measured with systematic
variations in wind speed, size of turbulence produc-
ing grid, and downstream location of heat source.
The nature of the temperature fluctuation field has
been studied. A comparison of Lagrangian and
Eulerian analyses for diffusion in a nondecaying tur-
bulence yields an expression for turbulent-heat-
transfer coefficient in terms of turbulence velocity
and a Lagrangian "scale." A convenient form has
been deduced for the criterion of interchangeability of
instantaneous space and time derivatives in a flowing
turbulence.

NACA TN 2712

FLOW CHARACTERISTICS OVER A LIFTING
WEDGE OF FINITE ASPECT RATIO WITH
ATTACHED AND DETACHED SHOCK WAVES AT A
MACH NUMBER OF 1.40. John H. Hilton, Jr. June
1952. 21p. diagrs., photos. (NACA TN 2712)

A series of schlieren photographs and pressure dis-
tributions obtained in the Langley 4- by 4-foot super-
sonic tunnel are presented to illustrate the general
characteristics of supersonic flow over a wedge
under conditions of attached and detached shock. The
pressure distributions and aerodynamic coefficients
are compared with shock-expansion theory.


5


NACA TN 2713

EFFECT OF COMPRESSOR-OUTLET AIR BLEED ON
PERFORMANCE OF A CENTRIFUGAL-FLOW TUR-
BOJET ENGINE WITH A CONSTANT-AREA JET
NOZZLE. Sidney C. Huntley. June 1952. 20p.
diagrs. (NACA TN 2713)

The effect of compressor-outlet air bleed on the per-
formance of a centrifugal-flow turbojet engine is
presented. The cost of compressor-outlet air bleed
in terms of the increase in net-thrust specific fuel
consumption was about 2 percent for each percent of
air bleed for constant engine speed, maximum net
thrust, or constant net-thrust operation. When maxi-
mum net-thrust operation with air bleed occurred at
a constant tail-pipe temperature, the maximum net
thrust decreased at a rate of about 2.5 percent for
each percent of air bleed, while the loss in net thrust
with air bleed at constant engine speed was only about
0.5 percent for each percent of air bleed.


NACA TN 2714

NORMAL ACCELERATIONS AND OPERATING CON-
DITIONS ENCOUNTERED BY A HELICOPTER IN
AIR-MAIL OPERATIONS. Almer D. Crim and Marlin
E. Hazen. June 1952. 17p. diagrs., photos., 2 tabs.
(NACA TN 2714)

An analysis is presented of the normal accelerations
and operating conditions encountered by a single-
rotor helicopter engaged in air-mail operation in the
vicinity of Los Angeles and its suburbs.


NACA TN 2716

EFFECT OF OPEN CIRCULAR HOLES ON TENSILE
STRENGTH AND ELONGATION OF SHEET SPECI-
MENS OF A MAGNESIUM ALLOY. R. S. Barker,
Aluminum Company of America. June 1952. 24p.
diagrs., 6 tabs. (NACA TN 2716)

The effect of open circular holes on the tensile
strength and elongation of sheet specimens of mag-
nesium alloy AM-C52S in both the annealed and the
hard-rolled condition was investigated. Tests were
made to study the effect of variable ratio of hole
diameter to total specimen width and also the effect
of spacing and arrangement of the holes. It was
found that greater reductions in strength were ex-
hibited by the annealed material but that the hard-
rolled material showed the greater reductions in
elongation.



NACA TN 2717

EFFECT OF TEMPERATURES FROM -70 T0600 F
ON STRENGTH OF ADHESIVE-BONDED LAP SHEAR
SPECIMENS OF CLAD 24S-T3 ALUMINUM ALLOY
AND OF COTTON- AND GLASS-FABRIC PLASTIC
LAMINATES. H. W. Eickner, W. Z. Olson and R. F.
Blomquist, Forest Products Laboratory. June 1952.
26p. diagrs., 6 tabs. (NACA TN 2717)








6


The performance of 14 commercial adhesives at tem-
peratures from -70 to 600 F was evaluated in lap
shear specimens of clad 24S-T3 aluminum alloy to
itself and that of 7 commercial adhesives at -70 to
250 F in lap joints of cotton-fabric-phenolic lami-
nate to itself, of glass-fabric-polyester laminate to
itself, and in joints of each of these laminates to clad
aluminum. One hot-setting tape adhesive was found
to be consistently superior to all others in lap-joint
specimens of aluminum tested at 450 F after 192
hours at 450 F. The best of the commercial adhe-
sives evaluated at -70 to 250 F in lap shear speci-
mens of plastic laminates bonded to themselves and to
aluminum had only fair resistance to stressing imme-
diately upon reaching 250 F. The adhesives general-
ly performed adequately in the various joints at -70 F.


NACA TN 2718

TWO-DIMENSIONAL STEADY NONVISCOUS AND
VISCOUS COMPRESSIBLE FLOW THROUGH A SYS-
TEM OF EQUIDISTANT BLADES. Hans J. Reissner,
Leonard Meyerhoff and Martin Bloom, Polytechnic
Institute of Brooklyn. June 1952. 48p. diagrs.,
4 tabs. (NACA TN 2718)

The problem of two-dimensional flow of a compressi-
ble nonviscous fluid through blade systems of equidis-
tant spacing, of identical shape, and of straight-line
arrangement of position is analyzed. Viscous flow
through a grid system of equidistant, narrowly spaced
blades is treated in an appendix.


NACA TN 2719

INVESTIGATION OF STATISTICAL NATURE OF
FATIGUE PROPERTIES. E. Epremian and R. F.
Mehl, Carnegie Institute of Technology. June 1952.
ii, 119p. diagrs., photos., tab. (NACA TN 2719)

Extensive fatigue tests were made on annealed Armco
iron and plain carbon and alloy steels heat-treated
to different strengths and microstructures. Statistics
of fatigue-fracture curves and endurance limits were
determined from the experimental data obtained and,
for various other materials, from a survey of litera-
ture. The results were analyzed to show the relative
effects of various metallurgical factors on the statis-
tical nature of fatigue properties. Other phases of
the problem studied include: dependence of statistical
variation in fatigue life on stress level in the fracture
range, statistics for location of crack initiation, size
effect, understressing effect, and the form and
method of plotting the S-N diagram.


NACA TN 2721

INITIAL RESULTS OF INSTRUMENT-FLYING
TRIALS CONDUCTED IN A SINGLE-ROTOR HELI-
COPTER. Almer D. Crim, John P. Reeder and James
B. Whitten. June 1952. 16p. diagrs., photos. (NACA
TN 2721)

Instrument-flying trials have been conducted in a
single-rotor helicopter to determine the adequacy of


NACA
RESEARCH ABSTRACTS NO.25


existing longitudinal flying-qualities requirements
under instrument conditions. In addition, lateral-
directional characteristics were examined. The suit-
ability, for helicopter use, of standard airplane in-
struments was also investigated.


NACA TN 2722

DISPLACEMENT EFFECT OF A THREE-
DIMENSIONAL BOUNDARY LAYER. Franklin K.
Moore. June 1952. 15p. diagrs. (NACA TN 2722)

A method is described for determining the "displace-
ment surface" of a known three-dimensional boundary-
layer flow in terms of the mass-flow defects asso-
ciated with the profiles of the two velocity compo-
nents parallel to the surface. The displacement sur-
face height is shown to differ, in general, from that
associated with the resultant mass-flow defect, even
at stagnation points of the secondary flow. Numeri-
cal values are found for the known three-dimensional
boundary-layer flow about a cone at a small angle of
attack to a supersonic stream.


NACA TN 2723

USE OF THE BOUNDARY LAYER OF A CONE TO
MEASURE SUPERSONIC FLOW INCLINATION.
Franklin K. Moore. June 1952. 21p. diagrs.
(NACA TN 2723)

An instrument is suggested for the measurement of
supersonic flow inclination, taking advantage of the
effect of angle of attack on the meridional velocity
profile of the boundary layer on a cone. This effect
of angle of attack may be measured by the difference
of total pressure recorded by two probes pointing
toward the apex and located in the plane of symmetry
of the flow.


NACA TN 2725

INTERACTION OF OBLIQUE SHOCK WAVES WITH
REGIONS OF VARIABLE PRESSURE, ENTROPY,
AND ENERGY. W. E. Moeckel. June 1952. 34p.
diagrs. (NACA TN 2725)

Equations are developed for computing the propaga-
tion of oblique shock waves through regions in which
pressure, energy, or entropy are continuously varia-
ble. For the special cases of supersonic shear flow
and Prandtl-Meyer flow, charts are given of shock
angle as function of Mach number for any initial-
strength wave. A sufficient condition for the avoid-
ance of upstream effects due to shock propagation Is
found. A wave-interaction procedure for determin-
ing supersonic portions of the flow downstream of the
shock is described, and an approximate method for
dealing with imbedded subsonic regions is proposed.






NACA
RESEARCH ABSTRACTS NO.25


NACA TN 2729

AN ANALYSIS OF SUPERSONIC FLOW IN THE RE-
GION OF THE LEADING EDGE OF CURVED AIR-
FOILS, INCLUDING CHARTS FOR DETERMINING
SURFACE-PRESSURE GRADIENT AND SHOCK-WAVE
CURVATURE. Samuel Kraus. June 1952. 45p.
diagrs., 5 tabs. (NACA TN 2729)

Flow in the region of the leading edge of curved air-
foils with attached shock waves is investigated. Ta-
bles and charts are presented for determining the
surface-pressure gradient and shock-wave curvature
in supersonic flow of an ideal diatomic gas. The re-
sults cover a range of Mach numbers from 1.5 to in-
finite and deflection angles from zero up to those ap-
proaching shock detachment. Calculations of surface-
pressure gradient and shock-wave curvature are also
made for curved airfoils in supersonic flow of a
calorically imperfect, diatomic gas. An approximate
procedure for determining the flow field a short dis-
tance downstream of the leading edge is also pre-
sented.


NACA TN 2731

INFLUENCE OF STRUCTURE ON PROPERTIES OF
SINTERED CHROMIUM CARBIDE. H. J. Hamjian
and W. G. Lidman. June 1952. 21p. diagrs., photos.,
5 tabs. (NACA TN 2731)

An investigation was conducted to study the influence
of structural variations on the properties of chromium
carbide sintered under pressure. The results show
that the room-temperature strength and hardness are
influenced by the stages of sintering, which are de-
fined by grain size and by the number, size, location,
and shape of pores. The extent to which sintering has
progressed during the second stage when densification
S occurs and the sintering conditions which will yield
optimum room-temperature strength can be deter-
mined from hardness. It was found that coarse-
grained structures are detrimental to room tempera-
ture strength. On the basis of limited data, coarse-
grained structures may not be detrimental at elevated
temperatures.


NACA RM 52D03

DESIGN OF APPARATUS FOR DETERMINING HEAT
TRANSFER AND FRICTIONAL PRESSURE DROP OF
NITRIC ACID FLOWING THROUGH A HEATED TUBE.
Bruce A. Reese and Robert W. Graham, Purdue Uni-
versity. June 1952. 61p. diagrs., photos., tab.
(NACA RM 52D03)

A design basis is presented for test apparatus to
measure heat-transfer and pressure-drop character-
istics of white fuming nitric acid under conditions
simulating those occurring in a regeneratively cooled
rocket engine. The apparatus is designed so that
heat-transfer and fluid-friction data can be obtained
over the following ranges: inlet pressure to test sec-
tion, atmospheric to 400 pounds per square inch ab-
solute; mean fluid temperature at entrance to heated
test section, -30 to 300 F; heat input to test section,
0.5 to 1.5 Btu per second per square inch; and
Reynolds number, 60,000 to 200,000.


7


NACA RM E52A11

ANALYSIS OF A PNEUMATIC PROBE FOR MEASUR-
ING EXHAUST-GAS TEMPERATURES WITH SOME
PRELIMINARY EXPERIMENTAL RESULTS. Marvin
D. Scadron. May 1952. 26p. diagrs., 4 tabs. (NACA
RM E52A11)

A pneumatic probe based on continuity of mass flow
through two restrictions separated by a cooling
chamber was constructed to measure gas temperature
at and beyond the limit of thermocouples. This probe
consisted of a subsonic flat-plate orifice for the first
restriction and a sonic-flow converging-diverging
nozzle for the second restriction. The effect of varia-
tion in gas constants on the calibration is examined for
common engine-exhaust gases. A high-temperature
wind tunnel that allowed calibration of the probe at
temperatures up to 2000 R and Mach numbers up to
0. 8 is described. Agreement to better than 30 R
between pneumatic probe indication and the indication
of a rake of radiation-shielded thermocouples indi-
cates that extrapolation of the calibration to higher
temperatures is possible with fair accuracy.


NACA RM E52C05

INVESTIGATION OF EFFECTIVE THERMAL CON-
DUCTIVITIES OF POWDERS. R. G. Deissler and
C. S. Eian. June 1952. 44p. diagrs., tab. (NACA
RM E52C05)

A simplified analysis was made to determine the
effective thermal conductivity of a powder from the
fraction of space occupied by the gas and the conduc-
tivities of the solid and the gas which make up the
powder. In order to check the analysis and to obtain
data of current technical interest, tests were con-
ducted to determine the conductivity of magnesium
oxide powder in various gases at temperatures
between 200 and 800 F. Good agreement was ob-
tained between analytical and experimental results.
The effects of some of the factors neglected in the
simplified analysis such as the bending of heat-flow
lines and the irregularity of the arrangement of the
particles were investigated and discussed.


NACA RM E52D03

PRELIMINARY SURVEY OF BOUNDARY-LAYER
DEVELOPMENT AT A NOMINAL MACH NUMBER OF
5.5. Harold L. Bloom. June 1952. 26p. diagrs.,
photos. (NACA RM E52D03)

Mean skin-friction coefficients on a flat-plate model,
with and without initial roughness, and on a wind tun-
nel wall were measured at a nominal Mach number of
5.5 over a Reynolds number range from I x 106 to
I x 107, and the results were compared with analyti-
cal values. Although evidence of air condensation was
obtained in the test section, experimental mean skin-
friction coefficients on the tunnel wall and on the flat
plate with artificial transition agreed quite well with
the analytical results of H. U. Eckert and E. R. Van
Driest. Experimental skin-friction coefficients on
the plate with natural transition fell between theoreti-








8


cal laminar values and the analytical turbulent values
of Eckert and of Van Driest. Because of the presence
of air condensation, the results reported herein must
be regarded as tentative.


NACA RM L52A28

NORMAL ACCELERATIONS AND ASSOCIATED
OPERATING CONDITIONS ON FOUR TYPES OF
COMMERCIAL TRANSPORT AIRPLANES FROM
VGH DATA AVAILABLE AS OF SEPTEMBER 1951.
Roy Steiner and Doris A. Persh. May 1952. 8p.
diagrs., 5 tabs. (NACA RM L52A28)

The NACA VGH time-history records of normal
acceleration, airspeed, and altitude of 377, 520, and
118 hours each for 3 four-engine aircraft and 377
hours for a twin-engine aircraft have been evaluated.
The summaries of gust loads and the associated
operating conditions are presented.


NACA TM 1331

INVESTIGATIONS OF THE BOUNDARY-LAYER
CONTROL OF A FULL SCALE SWEPT WING WITH
AIR BLED OFF FROM THE TURBOJET.
(Recherches sur l'Hypersustentation d'une Aile en
Fleche Reelle par Controle de la Couche Limite
Utilisant le Prel'evement d'Air sur le Turbo-
Reacteur). P. Rebuffet and Ph Poisson-Quinton.
April 1952. 43p. diagrs., photos. (NACA TM 1331.
Trans. from Recherche A6ronautique, no.14, March-
April, 1950, p.39-54)

The various stages of a boundary-layer control study,
culminating in the investigation of combined suction
and blowing in conjunction with a drooped-nose flap
and a form of double-slotted trailing-edge flap on a
full-scale 10-percent thick 31 20' sweptback wing
model in the Chalais-Meudon large-scale tunnel, are
described. The suction and blowing was accom-
plished with ejectors fed by bleed-off air from a
turbojet engine. A lift coefficient increment of 1.36
was obtained by the use of boundary-layer control.


NACA TM 1332

EXTENSION TO THE CASES OF TWO DIMENSIONAL
AND SPHERICALLY SYMMETRIC FLOWS OF TWO
PARTICULAR SOLUTIONS TO THE EQUATIONS OF
MOTION GOVERNING UNSTEADY FLOW IN A GAS.
(Estensione ai Cast di Simmetria Centrale Bi-e Tri-
Dimensionale di Due Particolari Soluzioni delle
Equazioni del Moto Gassoso Non Permanente).
Lorenzo Poggi. June 1952. 6p. (NACA TM 1332.
Trans. from Onore di Modesto Panetti, November 25,
1950.)

The author previously discovered two interesting
particular solutions to the equations of motion de-
scribing unsteady flow in a gas confined solely to a
one-dimensional duct. These solutions are now ex-
tended to cover the more noteworthy cases of central
symmetry in two and three dimensions.


iA

NACA -
RESEARCH ABSTRACTS NO.25


NACA TM 1334

THE EFFECT OF HIGH VISCOSITY ON THE FLOW
AROUND A CYLINDER AND AROUND A SPHERE.
(Der Einfluss grosser Zahigkeit bei der Stromung urn
den Zylinder und urn die Kugel). F. Homann. June
1952. 29p. diagrs., tab. (NACA TM 1334. Trans.
from Zeitschrift fur angewandte Mathematik und
Mechanik, v. 16, no. 3, June 1936, p. 153-164).

The report records an experimental and theoretical
investigation of the influence of high viscosity on the
stagnation pressure of flow around cylinders and
spheres. For the three-dimensional problem, a dif-
ferential equation is set up which corresponds to the
two-dimensional solution by Heimenz. Theoretical
methods are developed for determining the stagnation
pressure for the two-and the three-dimensional cases
of flow around cylinders and spheres, respectively.
The results thus obtained are compared with experi-
mental results obtained in an oil channel at Reynolds
numbers up to approximately 100. Finally, a proce-
dure to determine the velocity and the pressure varia-
tion, as well as the variation of the free-stream
velocity gradient on the stagnation streamline, is
shown and used for the practical case of Reynolds
number 100.


NACA TM 1338

THE OXIDATION OF METALS AND ALLOYS. (Uber
das Zundern von Metalle und Legierungen). Erich
Scheil. June 1952. 16p. diagrs., photos. (NACA
TM 1338. Trans. from Zeitschrift fur Metallkunde,
v. 29, July 1937, p. 209-214).

This paper reviews the various types of oxidation
processes occurring with pure metals and gives ex-
planations for the varying time-temperature-oxidation
rate relations that exist for copper, tungsten, zinc,
cadmium, and tantalum. The effect of shape and
crystal structure on oxidation is discussed. Princi-
ples derived are applied to the oxidation of alloys.


NACA TM 1339

VELOCITY OF ACTION OF OXYGEN, HYDROGEN
SULFIDE, AND HALOGENS ON METALS. (Die
Geschwindigkeit der Einwirkung von Sauerstoff,
Schwefelwasserstoff, und Halogenen auf Metalle).
G. Tammann and W. Koster. June 1952. 21p.
diagrs., 20 tabs. (NACA TM 1339. Trans. from
Zeitschrift fuir anorganische und allgemeine Chemie,
v. 123, August 1922, p. 196-201 and 208-224).

This report discusses a method of determining the
rate of surface oxidation of a metal by the change in
the color of the surface film produced by reactions
with oxygen, chlorine, or iodine. The metals studied
included iron, nickel, copper, zinc, cadmium, tin,
lead, cobalt, and manganese. Tables are given for
surface film thickness versus color for various
times.






NACA
RESEARCH ABSTRACTS NO. 25


NACA TM 1345


TRANSLATIONAL MOTION OF BODIES UNDER THE
FREE SURFACE OF A HEAVY FLUID OF FINITE
DEPTH. (0 postupatelnom dvizhenii tel pod svobod-
noi poverkhnost'yu tyazheloi zhidkosti konechnoi
glubiny). M. D. Haskind. June 1952. 20p. diagr.
(NACA TM 1345. Trans. from Prikladnaya Matemat-
ika i Mekhanika, v. 9, no. 1, September 1945,
p. 67-78).

The problem of the uniform motion of a solid body
under the free surface of a heavy Incompressible
fluid of finite depth is treated. The problem is a
two-dimensional one and is handled completely by
the methods of complex function theory.


N-3074B

INDEX OF NACA TECHNICAL PUBLICATIONS,
1949-MAY 1951. (Supplement to the index of NACA
technical publications, 1915-1949). viii, 201p.
(NACA)

This report, previously listed in Research Abstract
no. 24 as being available for loan only, is now avail-
able for issue.


BRITISH REPORTS


N-14296*

Aeronautical Research Council (Gt. Brit.)
GENERAL HANDLING TESTS OF THE SIKORSKY
R-4B HELICOPTER OVERFLYY MK. I). W. Stewart.
1951. 21p. diagrs., photos., tab. (ARC R & M 2431;
ARC 10,211. Formerly RAE Aero 2164)

Throughout the flying range, the stability of the heli-
copter is poor but the controls are very effective.
The lateral control may be criticized as too sensitive.
Maneuvers on the helicopter require a precise coordi-
nation of all controls, including the collective pitch
and throttle controls. Vibration in the control
column and lateral force required to trim, make it
very tiring to fly the Hoverfly for any length of time.
A loss of control is experienced at zero forward air-
speed and high power conditions, when the collective
pitch is reduced, unless the reduction is made very
slowly.


N-14304"

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF THE SWEEPBACK OF DELTA
WINGS ON THE PERFORMANCE OF AN AIRCRAFT
AT SUPERSONIC SPEEDS. A. Robinson and F. T.
Davies. 1951. 6p. diagrs. (ARC R& M 2476;
10,594. Formerly College of Aeronautics, Cranfield
Rept. 6)

The variation with sweepback of the total drag of an
aircraft in level flight at supersonic speeds is calcu-


9


lated. It is shown that sweepback is not uniformly
beneficial, but that in general the optimum amount of
sweepback depends on the design speed and altitude.


N-14305*

Aeronautical Research Council (Gt. Brit.)
INTERFERENCE ON A WING DUE TO A BODY AT
SUPERSONIC SPEEDS. S. Kirkby and A. Robinson.
1952. lOp. diagrs. (ARC R & M 2500; 10,631, FM
1113. Formerly College of Aeronautics, Cranfield
Rept. 7)

The wing-body combination considered consists of a
right-circular conical body, of small semivertical
angle, carrying an unswept symmetrical rectangular
wing whose chord is inclined at a small angle to the
axis of the body cone. In theory, the conical body is
replaced by an axial doublet system and the upwash
velocity increment is calculated at the midchord line
of the wing. The resultant lift increment on the wing
due to body interference is then calculated by
Ackeret's theory. Curves are plotted to show the
variation of induced wing-lift coefficient with cone
angle for different wing positions.


N-14306*

Aeronautical Research Council (Gt. Brit.)
OVERALL THRUST, THRUST GRADING, AND
TORQUE MEASUREMENTS ON A TWO-BLADE, 6-1/2
PER CENT THICK, NACA 16 SECTION PROPELLER
IN THE HIGH-SPEED TUNNEL. G. S. Hislop, G. F.
Hughes and D. S. Capps. 1951. 45p. diagrs., photos.,
tab. (ARC R & M 2515; 10,975. Formerly RAE
Aero 2213)

Blade angles were varied and the solidity was 7-1/2
percent total solidity at 0.7R. The Mach number
range was 0.1 to 0.8. It was concluded that it should
be possible to obtain very good propulsive efficiencies
at all forward speeds up to a Mach number of 0.78.
Beyond that speed, orthodox propellers will be unsuit-
able. There was some uncertainty about the correc-
tions to the experimental results.


N- 14307*

Aeronautical Research Council (Gt. Brit.)
HIGH-SPEED WIND-TUNNEL TESTS ON MODELS
OF FOUR SINGLE-ENGINED FIGHTERS (SPITFIRE,
SPITEFUL, ATTACKER AND MUSTANG). Staff of
the R.A.E. High Speed Wind Tunnel, edited by W. A.
Mair. PART I. W. A. Mair, S. P. Hutton and H. E.
Gamble. PART II. W. A. Mair, S. P. Hutton and
H. E. Gamble. PART IM, W. A. Mair and S. P.
Hutton. PART IV. S. P. Hutton, D. A. Clark and
D. J. Tremlett. PART V. J. Y. G. Evans,
J. Caldwell and C. M. Britland. 1951. 79p. diagrs.,
tabs. (ARC R & M 2535; ARC 6739; ARC 7596;
ARC 7045; ARC 9559; ARC 8707. Formerly RAE
Aero 1810; RAE Aero 1908; RAE Tech.Note Aero 1247;
RAE Aero 2112; RAE Aero 2038)

This report describes measurements of lift, drag,
and pitching moment made in the RAE high-speed








10


wind tunnel on models of the Spitfire, Spiteful
(F.l 1/43), Attacker (E.10/44), and Mustang. On the
Spiteful model, pressure distributions on the front
radiator flap were also measured. An introduction
(written in 1949) gives a general account of the tests
described in the separate parts of the report.


N-14308*

Aeronautical Research Council (Gt. Brit.)
TESTS ON A GLAS II WING WITHOUT SUCTION IN
THE COMPRESSED AIR WIND TUNNEL. C. Salter,
C. J. W. Miles and R. Owen. 1951. 21p. diagrs.,
5 tabs. (ARC R & M 2540. Formerly ARC 11,269;
Perf. 409; FM 1207)

The object of these tests was primarily the estimation
of the behavior of the wing at high Reynolds numbers
in the event of the failure of the suction, but the in-
vestigation was extended to the examination of the ef-
fect of main, split, and slotted flaps in order to ob-
tain information concerning some reasonable method
of countering any serious effects that might arise.
Critical regions were observed and the scale effects
were found to be large. The gain of maximum lift re-
sulting from the use of main and split flaps 6as rough-
ly 0.5 on CL and has been shown to be less than half
of the gain achieved on an NACA 0030 wing by the use
of a similar split flap. The effect of slotting the main
flap was found to be comparatively small and rather
detrimental than otherwise.


N-14309*

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE SOUTHWELL METHOD FOR ESTI-
MATING CRITICAL LOADS. H. L. Cox. 1951. 8p.
diagrs. (ARC R & M 2696. Formerly ARC 9852,
ARC 10,374, Strut 1054, Strut 1054a)

Draws attention to certain restrictions on the use of
the "Southwell plot" to estimate critical loads in
cases differing in conditions from those for which the
method was first proposed. The effect on this plot of
various stress distributions, of elastic failure of the
material and of other variations of critical stress,
however it may be occasioned, is examined. The
Southwell plot is strictly applicable only to deflections
which go to infinity at a definite critical load. In
other cases the plot usually overestimates the buck-
ling load, but the error should seldom be important.



N-14310*

Aeronautical Research Council (Gt. Brit.)
AN INVESTIGATION INTO THE EFFECT OF
FORCED AND NATURAL AFTERBODY VENTILA-
TION ON THE HYDRODYNAMIC CHARACTERISTICS
OF A SMALL FLYING BOAT (SARO 37) WITH A 1:20
FAIRING OVER THE MAIN STEP. J. A. Hamilton.
1951. 18p. diagrs., photos., tab. (ARC R & M 2714;
ARC 11,360. Formerly MAEE F/Res/206)


NACA
RESEARCH ABSTRACTS NO.25


A continuation of tests reported in R & M 2463 (in
which the fairing was 1:15). The hydrodynamic char-
acteristics were investigated in taxi runs and take-
offs over a range of fixed elevator positions and in
landing over a range of touch-down attitudes, at a
weight of 5,900 lbs. These tests were made with the
ducts naturally ventilated. At 6,200 lbs the tests were
confined to taxi runs with and without forced ventila-
tion.


N-14402*

Aeronautical Research Council (Gt. Brit.)
A REVIEW OF THE ESSENTIALS OF IMPACT FORCE
THEORIES FOR SEAPLANES AND SUGGESTIONS
FOR APPROXIMATE DESIGN FORMULAE. R. J.
Monaghan. 1952. 27p. diagrs. (ARC R & M 2720;
ARC 11,245. Formerly RAE Aero 2230)

This report reviews the essential theory and assump-
tions underlying recent work, and puts forward an
approximate design formula for the maximum deceler-
ation during a main step impact which is directly a
function of the initial impact conditions.


N-14403*

Aeronautical Research Council (Gt. Brit.)
A 24-WAY HIGH-SPEED ROTARY SWITCH FOR USE
IN STATIC AND AIRBORNE STRAIN GAUGE MEAS-
UREMENTS. D. H. Peirson. 1951. lOp. diagrs.,
photos. (ARC R & M 2232; ARC 10,398. Formerly
RAE Tech. Note SME 381)

A mechanical switch operating at 25 times per second
for each signal has been designed to record 24 strain-
gage signals, using a single amplifier and cathode-ray
oscillograph. This equipment has been used in the
measurement of steady ground and varying air-borne
signals. Reference is made to use of the switch in
static laboratory measurements on a tail plane and
wing. Technique as developed for the measurement
of wing stresses in flight is described in some detail.



N-14404*

Aeronautical Research Council (Gt. Brit.)
AN EXAMPLE IN WING THEORY AT SUPERSONIC
SPEED. H. B. Squire. 1951. 16p. diagrs., tab.
(ARC R & M 2549; ARC 10,624; ARC 12,517.
Formerly RAE Aero 2184)

Calculations of the pressure on a flat elliptic cone
and on a flat elliptic hyper-cone at supersonic speeds
and zero incidence are made for the case when the
cones lie inside the Mach cone of the apex. The re-
sults are combined to give the pressure distribution
and drag of a wing-like surface at zero incidence in a
supersonic stream. It is found that the pressure is
constant along straight lines on this surface which
are normal to the wind direction. The drag results
show the effect of sweepback on drag at supersonic
speeds.





NACA
RESEARCH ABSTRACTS NO. 25


N-14405*

Aeronautical Research Council (Gt. Brit.)
SOME DATA PERTAINING TO THE SUPERSONIC
AXIAL-FLOW COMPRESSOR. I. M. Davidson. 1951
38p. diagrs., photos., tab. (ARC R & M 2554; ARC
10,847; ARC 10,758. Formerly NOTE R.15; NGTE
Memo. M.16)

Together with some random considerations concerning
possible compressor development, data concerning
the flow of air at high speeds is presented in this note
in a form suitable for use in the design of supersonic
axial-flow compressors. A brief history and descrip-
tion is also given of the work of the German pioneers
Weise, Encke and Betz.


N-14406"

Aeronautical Research Council (Gt. Brit.)
TESTS ON A WORKING MODEL RAM JET IN A
SUPERSONIC WIND TUNNEL. J. R. Singham, F. W.
Pruden and R. C. Tomlinson. 1951. 16p. diagrs.,
photos., 4 tabs. (ARC R & M 2568; ARC 11,361.
Formerly NGTE R.20)

The tests were made to ascertain the practicability
of testing such a small-scale model in a wind tunnel
and to determine to what extent the external drag of a
model duct tested hot would differ from the same
model tested cold. The tests were run at a Mach
number of 1.4. It was concluded that satisfactory ac-
curacy could be obtained by the use of a large tunnel
of square or rectangular section, with a good velocity
distribution and transparent walls large enough to
view the whole model.


N-14409"

Aeronautical Research Council (Gt. Brit.)
DESIGNING TO AVOID DANGEROUS BEHAVIOUR OF
AN AIRCRAFT DUE TO THE EFFECTS ON CON-
TROL HINGE MOMENTS OF ICE ON THE LEADING
EDGE OF THE FIXED SURFACE. D. E. Morris.
1952. lOp. diagrs., tab. (ARC CP 66)

Results of wind-tunnel measurements of hinge mo-
ments of a Viking elevator with simulated ice on the
tail-plane leading edge are used to explain uncon-
trollable pitching oscillations of the aircraft which
occurred during a flight. It is suggested that the im-
posing of certain limitations on the control-surface
hinge-moment coefficients will eliminate the diffi-
culty.


N-14410*

Aeronautical Research Council (Gt. Brit.)
APPROXIMATE TWO-DIMENSIONAL AEROFOIL
THEORY. PART m. APPROXIMATE DESIGNS OF
SYMMETRICAL AEROFOILS FOR SPECIFIED PRES-
SURE DISTRIBUTIONS. S. Goldstein. 1952. 40p.
diagrs., 3 tabs. (ARC CP 70)

It is shown that if the velocity distribution at the sur-
face of a symmetrical airfoil, according to the purely


11



linear theory, is known, that the airfoil shape may be
easily calculated. General formulas are given for the
airfoil ordinates, both in terms of a Fourier series
and as an integral. Examples are given.


N-14411"

Aeronautical Research Council (Gt. Brit.)
FLIGHT TESTS ON THE YOUNGMAN-BAYNES
HIGH-LIFT EXPERIMENTAL AIRCRAFT. D. Lean.
1952. 54p. diagrs., photos., 9 tabs. (ARC CP 65)

The tests showed that an increment of maximum lift
coefficient of 1.32 can be obtained on an unswept wing
with a low-drag section, whose basic maximum lift
coefficient is 1.28. About 0.2 was lost due to wing-
fuselage interference. Adequate lateral control is
provided by ailerons inset in the full-span flap. The
profile-drag increment at full flap is 0.07 for a lift-
coefficient increment of 1.14, at an angle of attack of
10. Changes in longitudinal trim due to flaps are
small and easily controlled and the ground effect on
longitudinal trim is negligible. The wing-flap-aileron
structure is adequately stiff in torsion and the aileron
reversal speed is estimated at 300 knots.


N-14412*

Aeronautical Research Council (Gt. Brit.)
DRAFT CORRECTIONS FOR WATER SURFACE DE-
FLECTION UNDER NO. 1 CARRIAGE OF THE R.A.E.
SEAPLANE TANK. T. B. Owen. 1952. 15p. diagrs.,
photo., 3 tabs. (ARC CP 67)

The running draft and attitude of hulls under test are
measured relative to the undisturbed water level, but
at normal running speeds an appreciable deflection of
the water surface is produced at the model testing
position by the aerodynamic pressure field of the car-
riage. These deflections have been measured and
data are given for the correction of measured draft
and attitude values covering the normal range of
model positions and conditions.



N-14413'

Aeronautical Research Council (Gt. Brit.)
APPROXIMATE TWO-DIMENSIONAL AEROFOIL
THEORY. PART V. THE POSITIONS OF MAXI-
MUM VELOCITY AND THEORETICAL CL-RANGE.
S. Goldstein. 1952. 26p. diagrs. (ARC CP 72)

To calculate the theoretical critical compressibility
speed for a given airfoil section at a given CL we
first compute the greatest velocity on the airfoil con-
tour. Normal methods are often too long and thus a
simplification is presented. This method is accurate
enough for practical purposes. There is a discussion
of the theoretical CL-range for low-drag airfoils.
The so called "roof-top" airfoils are considered in
particular.








12


N-14414'

Aeronautical Research Council (Gt. Brit.)
THE MEASUREMENT OF HEAT TRANSFER AND
SKIN FRICTION AT SUPERSONIC SPEEDS.
PART U BOUNDARY LAYER MEASUREMENTS ON
A FLAT PLATE AT M = 2.5 AND ZERO HEAT
TRANSFER. R. J. Monaghan and J. E. Johnson.
1952. 44p. diagrs., photo., 2 tabs. (ARC CP 64)

The measured rate of growth of the laminar layer was
greater than theory would predict. Transition to tur-
bulent flow occurred between 3 and 4 inches from the
leading edge at a Reynolds number between 8 x 105
and 106. The thickness of the turbulent layer was
approximately the same as in low speed flow. Skin
friction was about 20 percent below that obtained with
the heated plate of the preceding report. A prelim-
inary investigation was made of the use of chemical
methods for indicating transition.



N-14415*

Aeronautical Research Council (Gt Brit.)
APPROXIMATE TWO-DIMENSIONAL AEROFOIL
THEORY. PART n. VELOCITY DISTRIBUTIONS
FOR CAMBERED AEROFOILS. S. Goldstein. 1952.
61p. diagrs., 5 tabs. (ARC CP 69)

Both thickness and camber are considered in this
theory. The connection with the "vortex sheet" theory
of infinitely thin airfoils is set out. In particular, the
usual formulas for the no-lift angle and moment at
zero lift are correct if account is taken of the thick-
ness, so long as the thickness is not too large. Ex-
amples are given.


N-14416 *

Aeronautical Research Council (Gt. Brit.)
APPROXIMATE TWO-DIMENSIONAL AEROFOIL
THEORY. PART I. VELOCITY DISTRIBUTIONS
FOR SYMMETRICAL AEROFOILS. S. Goldstein.
1952. 59p diagrs 7 tabs (ARC CP 68)

This part considers the velocity distribution at the
surface of a given airfoil. The airfoils are assumed
to have a thickness. Examples are given using the
NACA 0012, NACA 16-012, Clark Y, EQH 1260,
EQH 1250, and the EQH 1240 airfoils. The theory is
especially designed to be applied to airfoils with
ordinates given by algebraic formulas, or which can
be fitted to such formulas.


N-14417*

Aeronautical Research Council (Gt. Brit.)
APPROXIMATE TWO-DIMENSIONAL AEROFOIL
THEORY. PART IV. THE DESIGN OF CENTRE
LINES S Goldstein. 1952 49p. diagrs., 3 tabs.
(ARC CP 71)

This report is concerned with the design of the cen-
ter lines of cambered airfoils when a particular pres-
sure or velocity distribution is desired. Examples
are given.


NACA
RESEARCH ABSTRACTS NO.25


N-14652*

National Gas Turbine Establishment (Gt. Brit.)
STUDIES ON THE SPONTANEOUS IGNITION OF
FUELS INJECTED INTO A HOT AIR STREAM.
PART V. IGNITION DELAY MEASUREMENTS ON
HYDROCARBONS. B. P. Mullins. October 1951.
73p. diagrs., 46 tabs. (NGTE R.97)

Ignition delay measurements were made upon some
pure hydrocarbons and petroleum products using the
continuous method. The fuels were paraffins, aro-
matics, unsaturateds, naphthalene derivatives, cyclo-
compounds, terpenes, and petroleum products. Al-
though ignition delay curves of the petroleum prod-
ucts lie close together, considerable differences in
positions and slopes of the curves for single hydro-
carbon compounds were noted. Observations of flame
color were made and activation energies of reaction
were computed for each fuel.


N-14656"

National Gas Turbine Establishment (Gt. Brit.)
COMBUSTION OF DILUTE METHANE-AIR MIX-
TURES. H. Bruce Cox and N. P. W. Moore.
January 1952. 21p. diagrs., photo., 2 tabs. (NGTE
R. 113)

Combustion of these fuels has been investigated in a
conventional gas-turbine combustion chamber with a
secondary reaction duct at outlet temperatures be-
tween 650 and 900 C and from I to 3 atmospheres
static pressure. The efficiency reached 95 percent
at 890 C. At lower temperatures, the efficiency fell
due partly to the conversion to carbon monoxide of a
considerable part of the methane which did react.
Pressure has been shown to be unimportant within the
range investigated except for causing a small reduc-
tion in the yield of carbon monoxide at the higher
pressures. Oxidation in the secondary duct was slow.
Extended residence time was found to be essential for
high combustion efficiencies.


N-14706'

Royal Aircraft Establishment (Gt. Brit.)
METHOD OF QUALITATIVE BOUNDARY LAYER IN-
VESTIGATION BY MEANS OF HOT WIRES PLACED
IN THE WING BELOW STATIC PRESSURE ORIFICES.
(Methode voor kwalitatief grenslaagonderzoek met
behulp van gloeidraden zonder beinvloeding der
stroming). P. C. A. Malotaux, J. J. Denier van der
Gon and Yap Kie Jan. January 1952. 31p. diagrs.
(RAE Library Trans. 397. Trans. from Technische
Hoogeschool Le Delft, Sub-Aid. Vliegtuigbouwkunde,
Rapport VTH 45)

This report describes a method for qualitative bound-
ary layer investigation, using hot wires, in which the
flow remains undisturbed. Its application to the de-
termination of the local character of the boundary
layer flow past a wing surface is described. This can
be deduced from the behavior of hot wires located not
in tne boundary layer itself but below static pressure
orifices in the wing surface. As the flow remains
undisturbed the method can be combined with measure-
ments of the chordwise pressure distribution and
wake measurements of the profile drag.







NACA
RESEARCH ABSTRACTS NO. 25


N-14712*

Ministry of Supply (Gt. Brit.)
AN INVESTIGATION OF METHODS OF DUPLICA-
TION OF HYDRAULICALLY OPERATED IRREVERS-
IBLE FLYING CONTROLS. 132p. diagrs., photos.,
tab. (MOSS & TM 15 51; H. M. Hobson, Ltd.,
London)

The object of this report is to present an examination
of methods of duplicating hydraulically operated ir-
reversible flying controls of the piston and screw
jack types, with assessments of the relative merits of
the systems examined. Appendices to the report
describe tests of assemblies and components such as
hydraulic pumps, motors, piston jacks, screw jacks,
and control valves, arranged to represent certain of
the proposed methods.


N-14765*

Aeronautical Research Council (Gt Brit.)
USE OF THE PROPOSED ELECTRIC ISOGRAPH.
K. Mitchell. 1951. 24p. diagrs., 2 tabs. (ARC
R & M 2411; ARC 9014. Formerly RAE Tech. Note
Aero 1675)

This note examines in some detail the user's side of
the electric isograph proposed by Diprose. It is con-
cluded that a machine on these lines would be a most
valuable practical tool. Numerical methods for im-
proving approximate solutions determined by the iso-
graph are also given, and in particular a further in-
vestigation of Balrstow's method shows it to be suit-
able for determining complex roots of small modulus.
Where the modulus is large the method should be ap-
plied to the reciprocal equation.



N-14766*

Aeronautical Research Council (Gt. Brit.)
THE NO. 2 11-1/2 FT. x 8-1,'2 FT. WIND TUNNEL
AT THE ROYAL AIRCRAFT ESTABLISHMENT,
FARNBOROUGH. D. C. MacPhail, J. G. Ross and
E. C. Brown. 1951. 55p. diagrs., photos., 8 tabs.
(ARC R & M 2424; ARC 1945. Formerly RAE Tech.
Note Aero 1678)

Initially this fairly orthodox tunnel suffered from ob-
jectional flow characteristics common to many wind
tunnels. The causes of the troubles were investigated
and modifications recommended. These are de-
scribed. Some general conclusions are reached which
might be applicable to other wind tunnels.



N-14767 *

Aeronautical Research Council (Gt. Brit.)
FLIGHT TESTS ON THE PERFORMANCE OF
METEOR IV (A TWIN-ENGINED, SINGLE-SEAT JET
FIGHTER). F. Smith, D. J. Higton and R. H.
Plascott. 1951. lIp. diagrs., tab. (ARC R & M
2446; 10,190. Formerly RAE Aero 2148)


13


Tests were made to measure performance in level
flight and climb. Engine thrust was also measured
and nondimensional curves of net and gross thrust are
given. From the engine thrust and level flight results
the aircraft drag coefficient has been obtained up to a
Mach number of 0.81. The drag is somewhat less
than that calculated from dive tests on a Meteor I air-
craft.


N-14768"

Aeronautical Research Council (Gt. Brit.)
SOME EXPERIMENTS ON THE RESISTANCE OF
METALS TO FATIGUE UNDER COMBINED
STRESSES. H. J. Gough. PART I. H. J. Gough
and H. V. Pollard. PART II. H. J. Gough and W.
J. Clenshaw. 1951. 141p. diagrs., photos., 28 tabs.
(ARC R & M 2522; ARC 1397; 1873; 2575; 2684;
2685; 3168; 3703; 4041; 4270; 4585; 4755)

Presents information on fatigue under a combination
of reversed bending and reversed torsional stresses
(two independent variables). Next presents informa-
tion on fatigue strength of a nickel-chromium-
vanadium-molybdenum steel under a combination of
reversed bending stresses, reversed torsional
stresses, and superimposed static bending and tor-
sional stresses (four independent variables).


N-14769*

Aeronautical Research Council (Gt. Brit.)
SUCTION-SLOT DUCTING DESIGN. A. G. Rawcliffe.
1952. 14p. diagrs. (ARC R & M 2580; ARC 9487.
Formerly ARC 10,522; FM 1098; Perf. 304)

The ducting is to provide uniform suction through a
narrow slot along the span of a wing, with the lowest
possible losses, when the pump is situated at the root
of the wing. Models were tested and modified in the
light of results. A qualitative analysis is made and
recommendations have been formulated. Investiga-
tions were confined to still air.


N-14770*

Aeronautical Research Council (Gt. Brit.)
CALCULATION OF THE INFLUENCE OF A BODY
ON THE POSITION OF THE AERODYNAMIC CENTRE
OF AIRCRAFT WITH SWEPT-BACK WINGS.
H. Schlichting. 1952. 13p. diagrs., 3 tabs. (ARC
R & M 2582; ARC 10,689. Formerly RAE Tech.
Note Aero 1879)

From systematic three-component measurements of
wing-body combinations with swept wings it has been
found that the movement of the aerodynamic center
due to the influence of the body is greater for a swept-
forward than for a straight wing and less for a swept-
back wing. The forward shift of the aerodynamic
center due to the body for normal wing-body combi-
nations is about 0.6 chord for a straight wing, about
0.12 chord for a sweptforward, but about 0 for a 45
sweptback wing. A method of calculation is
presented.







14


N-14771*

Aeronautical Research Council (Gt. Brit.)
AN EXAMINATION OF THE TECHNIQUE OF THE
MEASUREMENT OF THE LONGITUDINAL
MANOEUVRING CHARACTERISTICS OF AN AERO-
PLANE, AND A PROPOSAL FOR A STANDARDISED
METHOD. D. J. Lyons. 1952. 33p. diagrs. (ARC
R & M 2597; ARC 11,159. Formerly RAE Aer6 2223)

It is demonstrated that the "steady stick force per g"
as defined by Gates and Lyon in R & M 2027 is the
best criterion for the measurement of maneuverability
of an aircraft because it indicates the minimum stick
force that has to be exerted by the pilot to break the
aircraft and its value is obtained in flight by a defi-
nite test procedure. It is further concluded that some
additional criterion may be used to insure that unduly
heavy forces are not encountered during sharp pull-
outs. A method of measuring the steady stick force
per g has been developed at the RAE which, it is
suggested, should be standardized for such tests
throughout the country. The results of this method
have been demonstrated on two aircraft.


N-14773*

Aeronautical Research Council (Gt. Brit.)
PRESSURE PLOTTING TESTS IN THE ROYAL AIR-
CRAFT ESTABLISHMENT HIGH SPEED WIND TUN-
NEL ON A 21 PER CENT THICK, LOW DRAG AERO-
FOIL (BRABAZON I WING ROOT SECTION). A. B.
Haines and W. Port. 1952. 24p. diagrs., 2 tabs.
(ARC R & M 2617; ARC 11,116. Formerly RAE Aero
2224)

The subject tests covered a Mach number range up to
0.7 at a Reynolds number of about 3 x 106 and low
speed tests were extended up to R = 9. 45 x 106. The
results have shown that: (1) No serious compressi-
bility effects on CL and Cm occur in the cruising con-
ditions (M up to 0.6, CL up to 0.66). (2) CLmax re-
mains roughly constant at about 1.15 up to M = 0.6,
and then falls to 1.01 for M = 0.7. (3) For M =0.2,
CLmax increases from 1.17 for R = 3 x 106 to 1.26 for
max
R = 9.45 x 106. (4) The reason for the maintenance of
CLmax seems to be a backward movement of the peak
suction which is not found on other sections. (5) Two
distinct CL versus a curves beyond the stall are ob-
tained at low Mach number. It appears that there is
adequate margin between the cruising and stalling con-
ditions to provide maneuverability and safety in up-
gusts.



N-14865*

Signals Research and Development Establishment
(Gt. Brit.) THE SINE SPRING. E. R. Wigan.
February 1949. 35p. diagrs. (SRDE Rept.1029)

Part I deals with the basic differences which distin-
guish this sine spring from known types of springs,
and describes a few of the unique mechanisms which


NACA
RESEARCH ABSTRACTS NO.25


arise out of its special properties. Part UI describes
the experimental evidence upon which rests the de-
sign data presented.


N-14997*

Marine Aircraft Experimental Establishment
(Gt. Brit.) PICK-UP HOOK FOR MOORING FLYING
BOATS. February 1952. 3p. diagrs. (MAEE
FX/Eq/317)

This hook has been found to be easier to handle and
lighter and stronger than the present standard
"grabbit" hook used on flying boats. No significant
corrosion has been noted. The only adverse com-
ment is that the highly polished handle becomes slip-
pery when wet. This difficulty could be remedied
easily. The hook floats horizontally in water.


N-16002*

Royal Aircraft Establishment (Gt. Brit.)
A NULL METHOD FOR THE CALIBRATION OF
FOUR-ARM STRAIN GAUGE BRIDGES. K. R.
Maslen. February 1952. 12p. diagrs. (RAE Tech.
Note Instn.130)

Describes a method by which the differential change
of resistance of the four arms of a closed Wheatstone
bridge may be measured directly by the use of an
auxiliary bridge. Only one source of power is re-
quired, and the output is measured in terms of the
differential change of resistance in the auxiliary
bridge necessary for balance. The ratio of this
change of resistance to the change in the strain-gage
bridge is independent of the gage resistance, the gal-
vanometer resistance, and the applied voltage.



N-16003*

Royal Aircraft Establishment (Gt. Brit.)
INVESTIGATIONS ON THE PRONE POSITION FOR
PILOTS. (Versuche zur liegenden unterbringugn des
flugzeugfuhrers). L. Schmidt. January 1952. 25p.
diagrs., photos. (RAE Library Trans.396. Trans.
from Zentrale fur wissenschaftliches Berichtswesen
der Luftfahrtforschung, Berlin. UM 1297, May 13,
1944).

Tactical and physiological requirements necessitate
for certain purposes a prone position for pilots. To
investigate this arrangement and the associated prob-
lems of piloting, the Flugtechnische Fachgruppe Ber-
lin designed and constructed in its own workshop the
B.9 experimental aircraft. The present report re-
views the layout and design aspects and summarizes
the experiences gained during the flight testing of the
B.9.


N-16035*

Aeronautical Research Council (Gt. Brit.)
DESIGNING A SLOT FOR A GIVEN WALL VELOCI-
TY. A.Thom and Laura Klanfer. 1952. 12p.
diagrs., tab. (ARC CP 76)








NACA
RESEARCH ABSTRACTS NO. 25


Gives the results obtained arithmetically for the wall
shape of an expanding passage with specified constant
wall velocities. A slot is assumed to draw fluid from
the passage at the velocity discontinuity and the shape
of the slot entry is obtained. A cusp develops at the
Entrance to the slot and the effect of the remainder of
the field of rounding this cusp is considered in detail.


N-16037"

Aeronautical Research Council (Gt. Brit.)
HIGH SPEED WIND TUNNEL TESTS ON AN AIRFOIL
WITH AND WITHOUT TWO-DIMENSIONAL SPAN-
WISE BULGES. J. A. Beavan, E. W. E. Rogers and
R. Cartwright. 1952. 34p. diagrs., photos., 2 tabs.
(ARC CP 78)

One bulge had a height of 0.004 chord and extended
from 0.3-0.5 chord. The other bulge was one-half of
both these dimensions. Pressure distributions at low
speeds agreed with theory. The CL is generally
greater than for the plain airfoil, but the correspond-
ing changes in Cm are small except at higher
speeds. At low speeds and zero angle of attack,
boundary-layer transition took place near the bulge
centers compared with 0.75 chord for the plain air-
foil. The resulting increase in CD agrees with
theory. With increase in Mach number, an initial
rise in drag (due to shock waves on the bulge) occurs
at angles of attack of 0 and 2. This rise is halted
as the main shock wave and transition point move
back and the final drag rise takes place at about the
same Mach number as on the plain airfoil. Shock
waves are stronger and better defined on the warped
than on the plain airfoil.


N-16040*

Aeronautical Research Council (Gt. Brit.)
N.P.L. AEROFOIL CATALOGUE AND BIBLIOGRA-
PLY. R. C. Pankhurst. 1952. 20p. (ARC CP 81)

Catalogues airfoils which have been designed (or sub-
stantially modified) at the NPL and which have been
the subject of theoretical investigations, aircraft de-
sign studies, or wind-tunnel tests. Within these lim-
its, it is intended to be complete with minor excep-
tions. Each airfoil has been assigned an NPL num-
ber. The arrangement follows broadly the succes-
sive stages in the development of the design theory.
Each list is prefaced with a brief indication of the
family character of the airfoils it contains.



N-16041 *

Aeronautical Research Council (Gt. Brit.)
POWER REQUIREMENTS FOR DISTRIBUTED SUC-
TION FOR INCREASING MAXIMUM LIFT. R. C.
Pankhurst and N. Gregory. 1952. 7p. diagr. (ARC
CP 82)

Considers the power requirements for distributed
suction. It appears that they are low for take-off and
landing; no estimates can be made for the case of
high-speed maneuvers until tests have been made
under the conditions of compressible flow.


15



MISCELLANEOUS


NACA Rept. 1017

Errata No 1 on "INVESTIGATION OF FREQUENCY-
RESPONSE CHARACTERISTICS OF ENGINE SPEED
FOR A TYPICAL TURBINE-PROPELLER ENGINE".
Burt L Taylor, ID and Frank L Oppenheimer. 1951.



NACA Rept. 1036

Errata No. 1 on -EXPERIMENTAL INVESTIGATION
OF THE EFFECTS OF VISCOSITY ON THE DRAG
AND BASE PRESSURE OF BODIES OF REVOLUTION
AT A MACH NUMBER OF 1.5". Dean R. Chapman
and Edward W. Perkins. 1951.


NACA TN 2117

Errata No. 1 on "DESIGN AND APPLICATIONS OF
HOT-WIRE ANEMOMETERS FOR STEADY-STATE
MEASUREMENTS AT TRANSONIC AND SUPERSONIC
AIRSPEEDS". Herman H Lowell. July 1950


NACA TN 2285

Errata No. 1 on "DAMPING IN ROLL OF CRUCI-
FORM AND SOME RELATED DELTA WINGS AT
SUPERSONIC SPEEDS". Herbert S. Ribner.
February 1951.


NACA TN 2548

Errata No. 1 on "EQUAL-STRENGTH DESIGN OF
TENSION-FIELD WEBS AND UPRIGHTS". Ralph H.
Upson, George M. Phelps and Tung-Sheng Liu.
January 1952.


NACA TN 2575

Addendum No. 1 to "A FLIGHT INVESTIGATION OF
THE EFFECT OF CENTER OF GRAVITY LOCA-
TION ON GUST LOADS". Jack Funk and Earle T.
Binckley. December 1951.


NACA TN 2614

Errata No. 1 on "ANALYTICAL INVESTIGATION OF
SOME THREE-DIMENSIONAL FLOW PROBLEMS IN
TURBOMACHINES" Frank E. Marble and Irving
Michelson, California Institute of Technology. March
1952.


NACA TM 1286

Errata No. 1 on "METHOD OF SUCCESSIVE AP-
PROXIMATIONS FOR THE SOLUTION OF CERTAIN
PROBLEMS IN AERODYNAMICS". M. E. Shvets
April 1951.






NACA
16 RESEARCH ABSTRACTS NO.
University of Florida ALF

UNPUBLISHED PAPERS I11111 flllliii IlIIIIU

31262090797514
N-15019*

Massachusetts Inst. of Tech.
THE EIGENVALUES OF SOME STABILITY PROB-
LEMS INVOLVING VISCOSITY. Cathleen S.
Morawetz. December 1951. 37p. diagrs.
(Massachusetts Inst. of Tech.)

Small disturbances are assumed to occur in viscous
flow between moving walls, with a symmetrical
velocity profile between two fixed walls, and in a
boundary layer along a flat plate. Corresponding to
each such problem for the viscous fluid, there is a
simpler problem with viscosity zero. It is natural
to conjecture that, for large Reynolds numbers, the
eigenvalues in the complete viscous problem are re-
lated to those in the corresponding inviscid problem.
The present investigation justifies the relationship
between viscous and inviscid eigenvalues in most
cases and in fact expresses these viscous eigenvalues
asymptotically with respect to the Reynolds number in
terms of the inviscid eigenvalues.


DECLASSIFIED NACA REPORTS


NACA RM E8120
National Advisory Committee for Aeronautics.
COMPARISON OF NATIONAL BUREAU OF STAND-
ARDS CERAMIC COATINGS L-7C AND A-417 ON
TURBINE BLADES IN A TURBOJET ENGINE.
C. Robert Morse. December 22, 1948. photos.,
5 tabs. (NACA RM E8120) (Declassified from
Restricted, 3/28/52)

An investigation was conducted to determine which of
two ceramic coatings L-7C and A-417 developed by
the National Bureau of Standards was the more suit-
able as a protective coating for gas-turbine blades.
Four cast Vitallium blades, two coated with each of
these ceramics, were installed in the turbine wheel
of a turbojet engine and subjected to 20 minute cycles
consisting of 5 minutes at idle and 15 minutes at
rated speed. Ceramic coating A-417 showed no evi-
dence of fusion during 100 operating cycles. Ceramic
coating L-7C showed evidence of fusion and radial
flow, indicating that the fusion temperature of the
coating is too low for satisfactory service in the
turbojet engine.


NACA-Langley 7-1-52 4000




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