Research abstracts

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Title:
Research abstracts
Physical Description:
93 v. : ; 27 cm.
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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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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).

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University of Florida
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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:
AA00009235:00050

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National Advisory Committee for Aeronautics


Research Abstracts


NO. 86


JULY 18, 1955


CURRENT NACA REPORTS


NACA Rept. 1174

THE STRUCTURE OF TURBULENCE IN FULLY .-
DEVELOPED PIPE FLOW. John Laufer, onajl" '
Bureau of Standards. 1954. ii, 18p. diagr. '
(NACA Rept. 1174. Formerly TN 2954)
Measurements, principally with a hot-wire a emn-
eter, were made in fully developed turbulent low in
a 10-inch pipe at speeds of 10 and 100 feet pe -,. ..
second. It is shown that rates of turbulent-energy ,-
production, dissipation, and diffusion have sha p
maximums near the edge of the laminar sublaye and
that there exist a strong movement of kinetic energy
away from this point and an equally strong movement
of pressure energy toward it. It is suggested that the
flow field may be divided into three regions: Wall
proximity where turbulence production, transfer,
and viscous action are of about equal importance; the
central region of the pipe where energy diffusion
predominates; and the intermediate region where the
local rate of change of turbulent-energy production
dominates the energy received by diffusive action.


.t. I,


NACA Rept. 1181


STRUCTURAL RESPONSE TO DISCRETE AND
CONTINUOUS GUSTS OF AN AIRPLANE HAVING ,
WING-BENDING FLEXIBILITY AND A CORREL'XVLL
TION OF CALCULATED AND FLIGHT REStBITS.
John C. Houbolt and Eldon E. Kordes. 194. ,ii,
22p. diagrs., 4 tabs. (NACA Rept. 1181. Formerly
TN 3006) \

An analysis is made of the structural response to,-...._
gusts of an airplane having the degrees of freedom
of vertical motion and wing bending flexibility. Con-
venient solutions of the response equations are
developed for discrete and sinusoidal-gust encounter,
and the procedure is given for treating the realistic
condition of continuous random atmospheric turbu-
_.-.._-_L e-L .- ght and calculated results are then given
for several airplanes to evaluate the influence of
wing bending flexibility on the structural response to
gusts. The discrete-gust approach is shown to
reveal the general nature of the flexibility effects
and leads to qualitative correlation with flight re-
sults. The continuous-turbulence approach shows
good quantitative correlation and indicates a much
greater degree of resolution of flexibility effects.

NACA Rept. 1193

THEORETICAL PERFORMANCE CHARACTERIS-
TICS OF SHARP-LIP INLETS AT SUBSONIC
SPEEDS. Evan A. Fradenburgh and DeMarquls D.
Wyatt. 1954. ii, 8p. diagrs. (NACA Rept. 1193.
Formerly TN 3004)


A method is presented for the estimation of the
subsonic-flight-speed characteristics of sharp-lip
inlets applicable to supersonic aircraft. The
analysis, based on a simple momentum balance con-
sideration, permits the computation of inlet pres-
Ssure recovery mass-flow relations and additive-
Sdrag coefficients for forward velocities from zero to
the speed of sound. Operation of a sharp-lip inlet at
velocity ratios less than 1.0 results in an additive
drag that is not cancelled by lip suction, while at
Velocity ratios greater than 1.0, losses in inlet total
pressures result In particular, at the take-off con-
di lon, -he toalpressure and the mass flow for a
-chokedde-W are only 79 percent of the values
--deally attainable with a rounded lip.

NACA RM 54L29

METEOROLOGICAL PROBLEMS ASSOCIATED
WITH COMMERCIAL TUBOJET-AIRCRAFT OPER-
ATION. A working group of the NACA Subcommittee
on Meteorological Problems. June 1955. 46p.
(NACA RM 54L29)

Meteorological requirements and problems antici-
pated with the safe, efficient, and economical oper-
ation of commercial turbojet aircraft have been an-
alyzed and evaluated. Problems of temperature,
c"wind, pressure, ceiling, visibility, cloud and clear-
r turbulence, icing, and communication are dis-
sed. Characteristics such as high cruising speed,
cruise level, relatively high fuel consumption,
engine performance sensitive to temperature and
55 ar density are considered.

ACA RM E55D07b

EXPLORATORY INVESTIGATION OF FLOW IN THE
SEPARATED REGION AHEAD OF TWO BLUNT
BODIES AT MACH NUMBER 2. Harry Bernstein
and William E. Brunk. June 1955. 27p. diagrs.,
photos. (NACA RM E55D07b)

Flow separation from a flat plate ahead of two blunt
two-dimensional bodies was investigated at Mach
number 2.0. Interferograms were obtained for the
regions, and the pitot-pressure distribution and
flow directions were surveyed in one of the regions.
Mach numbers were generally less than 0.5 in the
reverse-flow regions near the plate surface. A
flow-direction survey for a model with separation at
the leading edge of the flat plate showed a reverse-
flow component in about half of the separated region.

NACA RM L55D21

VELOCITY DISTRIBUTIONS MEASURED IN THE
SLIPSTREAM OF EIGHT-BLADE AND SIX-BLADE
DUAL-ROTATING PROPELLERS AT ZERO
ADVANCE. Leland B. Salters, Jr. June 1955.
27p. diagrs., photo. (NACA RM L55D21)


*AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW., WASHINGTON 25, D C., CITING CODE NUMBER ABOVE EACH TITLL.
THE REPORT TITLE AND AUTHOR.


fz? / ;72-





NACA
RESEARCH ABSTRACTS NO. 86


This report contains the results of a slipstream
survey behind NACA 8.75-(5)(05)-037 eight- and six-
blade dual-rotating propellers at zero advance. The
slipstream-boundary cone angle was found to agree
qualitatively with the theoretical angle of spread for
an unheated jet.

NACA RM L55E10b

EXPERIMENTAL ANALYSIS OF MULTICELL
WINGS BY MEANS OF PLASTIC MODELS. George
W. Zender. June 1955. 6p. diagrs. (NACA
RM L55E10b)

The stresses and deflections of a plastic model of a
delta multicell wing are compared with results
obtained by the use of the Cal-Tech analog computer.
The comparison indicates that valuable information
may be obtained for experimental structural analyses
from tests of plastic models.

NACA RM L55E12c

VERTICAL AND DRAG GROUND-REACTION
FORCES DEVELOPED IN LANDING IMPACTS OF A
LARGE AIRPLANE. Richard H. Sawyer, Albert W.
Hall and James M. McKay. June 1955. 12p. diagrs.
(NACA RM L55E12c)

Tests were conducted on a large-bomber-type air-
plane to determine the ground reactions imposed on
the landing gear under actual landing conditions.
The program covered landings made at vertical
velocities up to 8.5 feet per second and forward
speeds at contact from 95 to 120 miles per hour.
Landings were made on both wet and dry concrete
runways. Results are presented of the effects of
vertical velocity at contact and the effects of runway
surface condition (wet and dry) on the vertical and
drag ground reactions obtained during the landing
impact.

NACA TM 1391

REDUCTION OF THE SHIMMY TENDENCY OF TAIL
AND NOSE-WHEEL LANDING GEARS BY INSTAL-
LATION OF SPECIALLY DESIGNED TIRES.
(Verminderung der Flatterneigung von Sporn- und
Bugwerken durch Einbau besonders geformter
Reifen). H. Schrode. July 1955. 13p. diagrs.
(NACA TM 1391. Trans. from Deutschen
Versuchsanstalt fiir Luftfahrt E. V., Berlin-
Adlershof. Technische Berichte, v. 10, 1943,
p. 113-116)

An experimental study is made of the shimmy
tendency of several conventional and modified
German aircraft tires ranging in size from about
11 to 15 inches in diameter. The effects of tire
size, shape, loading and wear, type of rolling mo-
tion (acceleration or deceleration), trail and rolling
velocity on the shimmy tendency are investigated.
It is found that tires of small main dimensions with
stiff flattened running surfaces, high lateral and
torsional stiffnesses and small ground contact areas
are favorable for reduction of the shimmy tendency.

NACA TN 3299

MAXIMUM THEOREMS AND REFLECTIONS OF
SIMPLE WAVES. P. Germain, Brown University.
.June 1955. 22p. (NACA TN 3299)

The properties of the solutions corresponding to the
reflection of a centered simple wave along a straight


line or along a free streamline are shown to be
closely related to some important theorems predict-
ing "a priori bounds" for special mathematical
problems. These properties thus appear to be a
physical interpretation of those theorems.

NACA TN 3421

AERODYNAMICS OF A RECTANGULAR WING OF
INFINITE ASPECT RATIO AT HIGH ANGLES OF
ATTACK AND SUPERSONIC SPEEDS. John C.
Martin and Frank S. Malvestuto, Jr. July 1955.
114p. diagrs., tab. (NACA TN 3421)

Perturbation of the flow over a two-dimensional flat
plate at finite angles of attack is used to obtain a first-
order evaluation of damping in roll, lift and moment
due to an increment in angle of attack, and lift and
moment due to steady pitching velocity for a rectan-
gular wing of infinite aspect ratio at supersonic
speeds. Results are valid for the range of Mach
number and angle of attack for which the flow behind
the shock is supersonic. The analysis is based on
the equations for rotational flow so that the change in
entropy is taken into account. Approximate estunates
of a number of aerodynamic derivatives of rectan-
gular wings at finite angles of attack are also pre-
sented.

NACA TN 3443

SHEARING EFFECTIVENESS OF INTEGRAL
STIFFENING. Robert F. Crawford and Charles
Libove. June 1955. 37p. diagrs., photo., tab.
(NACA TN 3443)

Values of coefficients for defining the effectiveness
of integral stiffeners in resisting shear deformations
of the plate of which they are an integral part are
presented for a variety of proportions of rectangular
stiffeners with circular fillets. Formulas are given
in which these coefficients may be employed to cal-
culate the elastic constants associated with the twist-
ing and shearing of integrally stiffened plates. The
size of fillet radius is shown to contribute appreciab-
ly to the degree of penetration of the stresses from
the skin into the stiffener.

NACA TN 3455

RECOVERY AND TIME-RESPONSE CHARACTER-
ISTICS OF SIX THERMOCOUPLE PROBES IN SUB-
SONIC AND SUPERSONIC FLOW. Truman M.
Stickney. July 1955. 25p. diagrs., photos., 2 tabs.
(NACA TN 3455)

Experimental data obtained from three shielded and
three unshielded thermocouple probes are presented.
Data taken in air at room temperature over the
ranges 0.2 to 2.2 Mach number and 0.2 to 2.2 atmos-
pheres total pressure show reproducible systematic
variations of recovery with Mach number, ambient
pressure, flow angle, and probe design. Time-
constant data determined at Mach 0.2 and room
temperature and pressure indicate that unshielded
probes are several times faster in response to
temperature changes than shielded probes.

NACA TN 3456

PROPAGATION OF A FREE FLAME IN A TURBU-
LENT GAS STREAM. William R. Mickelsen and
Norman E. Ernstein. July 1955. 89p. diagrs.,
photos., 2 tabs. (NACA TN 3456)





NACA
RESEARCH ABSTRACTS NO. 86


Effective turbulent free-flame speeds measured in
turbulent, flowing propane-air mixtures were found
to have statistical distributions about mean values.
The statistical spread was greater for rich and lean
fuel-air ratios and at high turbulence intensities.
The measured flame speeds, together with hot-wire-
anemometer measurements, formed a basis for com-
parison with three theories and other types of flames.
Although the free-flame speeds are lower than those
for turbulent Bunsen and stabilized flames, values
calculated from the Tucker analysis and a modified
Scurlock-Grover analysis seem to form an upper
limit to the turbulent free-flame-speed data.




NACA TN 3458

UNSTABLE CONVECTION IN VERTICAL CHANNELS
WITH HEATING FROM BELOW, INCLUDING EF-
FECTS OF HEAT SOURCES AND FRICTIONAL
HEATING. Simon Ostrach. July 1955. 38p. diagrs.,
3 tabs. (NACA TN 3458)

Solutions are found for the cases in which the wall
temperature variations are linear and (1) the wall
temperatures are specified, (2) the walls are both
insulated, and (3) the net mass flow in the channel
is zero. The heat sources affect the flows in es-
sentially a quantitative manner, changing the stability
characteristics of the flows only when both walls are
insulated. The effects of frictional heating are
important in certain ranges of the parametric values,
especially near critical Rayleigh numbers.




NACA TN 3459

SIMPLIFIED PROCEDURES AND CHARTS FOR THE
RAPID ESTIMATION OF BENDING FREQUENCIES
OF ROTATING BEAMS. Robert T. Yntema. June
1955. ii, 90p. diagrs., 6 tabs. (NACA TN 3459.
Supersedes and extends RM L54G02)

A Rayleigh energy approach utilizing the nonrotating-
beam bending modes in the determination of the bend-
ing frequencies of the rotating beam is evaluated and
is found to give good practical results for helicopter
blades. Charts are presented for the rapid estima-
tion of the first three bendig frequencies for rotat-
ing and nonrotating cantilever and hinged beams. A
more exact mode-expansion method used in evaluat-
ing the Rayleigh approach is also described. Nu-
merous mode shapes and derivatives are presented
in tabular form and discussed.




NACA TN 3500

CORRECTION OF ADDITIONAL SPAN LOADINGS
COMPUTED BY THE WEISSINGER SEVEN-POINT
METHOD FOR MODERATELY TAPERED WINGS
OF HIGH ASPECT RATIO. John DeYoung and
Walter H. Barling, Jr. July 1955. 31p. diagrs.
(NACA TN 3500)

A simple procedure is found which results in more
accurate span loadings, lilt-curve slopes, and span-
wise centers of pressure being read directly from
the charts of NACA Report 921. The new results


compare very well with experiment and with theoret-
ical results believed 10 be accurate.



NACA TN 3501

THE TRANSONIC CHARACTERISTICS OF 22
RECTANGULAR. SYMMETRICAL WING MODELS
OF VARYING ASPECT RATIO AND THICKNESS.
Warren H. Nelson and John B. McDevitt. June 1955.
109p. diagrs., photos. (NACA TN 3501. Formerly
RM A51AI2)

An investigation utilizing the transonic-bump tech-
ruque was made to determine the aerodynamic char-
acteristics at transonic Mach numbers of 22 rec-
tangular wings having aspect ratios of 6. 4, 3, 2,
1.5, 1, and 0.5, and NACA 63AOXX sections with
thickness-to-chord ratios of 10, 8, 6, 4, and 2 per-
cent. The Mach number range was 0.4 to 1.1, cor-
responding under the test conditions to a Reynolds
number range from 1.25 to 2.05 million. These
data are presented without analysis.



NACA TN 3504

EFFECT OF TRAILING-EDGE THICKNESS ON LIFT
AT SUPERSONIC VELOCITIES. Dean R. Chapman
and Robert H. Kester. June 1955. 19p. diagrs.
(NACA TN 3504. Formerly RM A52D17)

Lift forces on various rectangular-plan-form wings
were measured in the Mach number range between
1.5 and 3.1 at Reynolds numbers between 0.55 and
2.2 million. The wings differed in trailing-edge
thickness, profile shape, maximum thickness ratio,
and aspect ratio. Measurements were made on wings
with and without a boundary-layer trip and are com-
pared to theoretical calculations. Calculated results
using shock-expansion theory are presented for Mach
numbers up to 10. In general, thickening the trailing
edge resulted in an increase in lift-curve slope. This
increase varied between a few percent and about 15
percent, depending primarily on the trailing-edge
thickness. Calculations indicate that somewhat
greater increases are possible at high supersonic
Mach numbers.


NACA TN 3508

LAMINAR FREE CONVECTION ON A VERTICAL
PLATE WITH PRESCRIBED NONUNIFORM WALL
HEAT FLUX OR PRESCRIBED NONUNIFORM WALL
TEMPERATURE. E. M. Sparrow. July 1955. 34p.
diagrs. (NACA TN 3508)

An analysts is made for laminar free convection on
a vertical plate with prescribed nonuniform thermal
conditions at the surface. For the situation where
the wall-heat-flux variation Is prescribed, graphs
are presented from which the.resulting variation of
the wall temperature and Local heat-transfer coef-
ficient can be found. Results for the important
special case of uniform wall heat flux are given.
For the situation where the wall-temperature varia-
tion is prescribed, graphs are given for the over-all
heat-transfer rate and local heat-transfer coefficient.
All results are given for Prandtl numbers from 0.01
to 1000. Boundary-layer theory is used in con-
junction with the KIrmin-Pohlhausen method.






NACA
RESEARCH ABSTRACTS NO. 86


NACA TN 3513

HEAT TRANSFER AT THE FORWARD STAGNATION
POINT OF BLUNT BODIES. Eli Reshotko and
Clarence B. Cohen. July 1955. 17p. diagrs.
(NACA TN 3513)

Relations are presented for the calculation of heat
transfer at the forward stagnation point of both two-
dimensional and axially symmetric blunt bodies.
The relations for the heat transfer, which were ob-
tained from exact solutions to the equations of the
laminar boundary layer, are presented in terms of
the local velocity gradient at the stagnation point.
These exact solutions include effects of variation
of fluid properties, Prandtl number, and transpira-
tion cooling. Examples illustrating the calculation
procedure are presented.


NACA TN 3517

APPROXIMATE METHOD FOR DETERMINING
EQUILIBRIUM OPERATION OF COMPRESSOR
COMPONENT OF TURBOJET ENGINE. Merle C.
Huppert. July 1955. 25p. diagrs. (NACA TN 3517)

A method is presented for estimating the equilibrium
operating line for a compressor as a component part
of a turbojet engine. The results of this analysis are
presented in chart form to facilitate rapid determina-
tion of the equilibrium operating line. Predictedand
measured equilibrium operating conditions are com-
pared in terms of compressor pressure ratio and
equivalent weight flow; satisfactory agreement is
indicated.




BRITISH REPORTS



N-38032*

Royal Aircraft Establishment (Gt. Brit.)
A REVISED ESTIMATE OF A RELATIONSHIP BE-
TWEEN BEARING STRENGTH AND HARDNESS FOR
METALLIC STRUCTURAL MATERIALS. R. F.
Mousley. January 1955. 9p. diagr., 4 tabs.
(RAE Tech. Note Structures 144)

A revised estimate is made of a relationship between
bearing strength and hardness for metallic structural
materials.


N-38033*

Royal Aircraft Establishment (Gt. Brit.)
THE EFFECT OF NUMBER OF RIGGING LINES ON
THE STRENGTH OF AN EXETER TYPE 12 PARA-
CHUTE CANOPY. J. Picken. January 1955. lip.
diagrs., photos., 2 tabs. (RAE Tech. Note Mech.
Eng. 197)

The number of rigging lines has been shown to have
an effect on the strength of an Exeter-type L2 para-
chute canopy. In a series of flight tests it was
demonstrated that a canopy rigged with eight lines
is not so strong as one rigged with 16 lines.
Amendments to the appropriate design formula are
suggested.


N-38034*

Royal Aircrait Estaolishment (Gt. Brit.)
THE ELECTRICAL RESIST VTY OF SODIUM BE-
TWEEN 780 K AND 3720 K. F. J. Bradshaw and
S. Pearson. February 1955. 14p. diagrs. (RAE
Met. 84)
The electrical resistivity of sodium has been meas-
ured from 780 K to slightly beyond the mellng point.
The specimens used consisted of distilled sodium in
fine thin-walled nickel tunes. The results are com-
pared with previously published figures. There is a
small rise in the R T v. T curve in the region 3000
K to the melting point. which is not simply account-
able for by present conductivity theories, and the
hypothesis that part of the high temperature resis-
tance is due to the presence of lattice defects is
considered. Measurements on solid sodium were
made to within 0.040 of the melting point and the
resistance increased uniformly to this temperature.
Some attempts *'ere made to measure excess resis-
tivity due to defects retained after quenching the
solid from temperatures near the melting point, or
after mechanical deformation at 630 K, but these
were unsuccessful.


N-38035*

Royal Aircraft Establishment (Gt. Brit.)
TITANIUM-VANADIUM ALLOYS. G. I. Lewis,
K. S. Jepson and H. Brooks. February 1955. 40p.
diagrs., photos., 12 tabs. (RAE Tech.NoteMet.208)

Titanium-base alloys containing 5- to 30-percent
vanadium were prepared as 0.07-in. thick rolled
strip and tested in various conditions to determine
the best tensile properties obtainable from them by
heat treatment. An alloy with 16-percent vanadium
showed most promise as a structural sheet material.
After quenching from 8000 C it had a tensile strength
of 51 tons/sq in. with 29-percent total and 26-
percent uniform elongation, and a minimum bend
radius of IT. By averaging at 400 to 4750 C, this
soft condition could be converted to various hard
conditions with strengths ranging from 71 tons/sq in.
with 12-1/2 percent elongation to 88 tons.'sq in. with
5-percent elongation. When aged to a 200 C
strength of 81 tons sq in., the tensile and shear
moduli were 14.3 and 5.4 x 106 lb, sq in., respec-
tively; the strength at 3500 C was 65 tons/sq in.
and after 1000 hours exposure at this temperature,
the 200 C properties were substantially unchanged.
The density of the alloy was 4.74 gm, cc or 0.17
lb/cu in. A number of ternary alloys based on the
Ti-16-percent V composition were also examined.


N-38041*

Royal Aircraft Establishment (Gt. Brit.)
A CONSTRUCTIONAL METHOD FOR MINIMISING
THE HAZARD OF CATASTROPHIC FAILURE IN A
PRESSURE-CABIN. D. Williams. March 1955.
7p. (RAE Tech. Note Structures 156)

A method is put forward for substantially reducing
the chances of a local failure in the shell of a
pressure cabin from developing into catastropicfail-
ure of the cabin. The increased safety is achieved
without weight penalty, and consists essentially in
using closely spaced (10 inches or thereabouts)
transverse flat bands, the material for which is
obtained by reducing the sheet thickness normally
available for the shell walls.







NACA
RESEARCH ABSTRACTS NO. 86


N-38042'

Nat. Gas Turbine Establishment (Gl. Brit.)
SOME EFFECTS OF SCALE AND MATERIAL ON
THE BURSTING SPEEDS OF TURBINE DISCS.
A. Graham, T. M. Jones and V. C. H. Bailey.
January 1955. 12p. diagrs., 5 tabs. (NGTE R.166)

The measured bursting speeds of turbine disks of
related profiles and several diameters in several
materials are compared. Within the experimental
range, the desirable amount of reinforcement about
a central bore is found to vary with material and
scale, but a moderate amount of reinforcement will
offset the weakness due to a bore. Larger disks in
various materials appear to behave like smaller
disks in a brittle material. It is concluded that
model experiments offer a better basis for compari-
son of disk profiles than elastic calculations.


N-38043'

Royal Aircraft Establishment (Gt. Brit.)
LUBRICATION: THE RATE OF SPREADING AND
CREEP OF OILS AND OTHER LIQUIDS OVER
SOLID SURFACES. E. B. Bielak, G. F. N.
Calderwood and E. J. W. Mardles. March 1955.
21p. diagrs., 3 tabs. (RAE Tech.Note Chem. 1245)

The rate of radial spreading of liquid pools over and
between horizontal surfaces and of the spontaneous
movement of liquid columns into horizontal capil-
laries has been measured for a variety of liquids
under different conditions of test. A resistance to
flow, due to thin films extending from the liquid
periphery, and which is variable with time of rest-
ing, rate of movement, etc., has been recorded.
The slow movement which is non-Newtonian and
affected by a number of specific factors other than
the physical ones of viscosity, density, etc. can be
correlated with results for several liquid properties,
that is, friction, adhesion, and the dispersive power
of the liquids for finely divided solids. The contact
angle and hysteresis of contact angles hypotheses
are inadequate for explaining the results obtained.
The significance of the results in relation to lubri-
cation theory is discussed briefly.


N-38044'

Royal Aircraft Establishment (Gt. Brit.)
LOADING CONDITIONS FOLLOWING AN AUTOMA-
TIC PILOT FAILURE (RUDDER CHANNEL). D. R.
Puttock. February 1955. 35p. diagrs., tab. (RAE
Tech. Note Structures 154)
A method is presented for the determination of the
critical loading conditions of aircraft that ensue
from an automatic pilot failure in the rudder channel.
General expressions have been derived through re-
sponse theory for the angle of sideslip, fin-and-
rudder load,and lateral acceleration both at the c. g.
of the aircraft and at the tail, that result from the
sequence of rudder movements assumed to follow an
automatic pilot failure. Analysis of these general
expressions leads to formulas suitable for assessing
the numercial values of the critical loading conditions
and it is suggested that these formulas might form a
basis for the interpretation of the appropriate design
requirement. An example is given to illustrate the
type of response produced by a rudder channel fail-
ure and the calculation procedure.


N-38045'
Royal Aircraft Establishment (Gt. Brit.)
LOADING CONDITIONS FOLLOWING AN AUTOMA-
TIC PILOT FAILURE (ELEVATOR CHANNEL).
D. R. Puttock. February 1955. 50p. diagrs., tab.
(RAE Tech. Note Structures 153)

A proposal is made for a standard procedure for
calculating the critical loading conditions ensuing
from an automatic pilot failure in the elevator chan-
nel. General expressions are derived through re-
spone theory for the increments in normal accelera-
tion at the c. g. of the aircraft, normal acceleration
at thetail, and aerodynamic load on the tailplane
which result from the sequence of elevator move-
ments assumed to follow a failure. Analysis of these
general expressions leads to formulas suitable for
assessing the numerical values of the critical loads
on the wing and tailplane. The influence of the se-
quence of elevator movements on the loading condi-
tions is discussed with reference to an example.



MISCELLANEOUS



N-37238'
INTERPRETATION OF WIND-TUNNEL DATA IN
TERMS OF DYNAMIC BEHAVIOR OF AIRCRAFT AT
HIGH ANGLES OF ATTACK. Ralph W. Stone, Jr.
(Presented to Wind Tunnel and Model Testing Panel
of Advisory Group for Aeronautical Research and
Development, Ottawa, Canada, June 10-14, 1955)
32p. diagrs.
With the advent of jet-propelled aircraft and flight
near and beyond the speed of sound, the basic airplane
configuration has changed markedly. Most notice-
able and important in this change has been the trend
to low-aspect-ratio wings, swept-wing plan forms,
and the general concentration of weight along the
fuselage. The aerodynamic characteristics which
basically result from the major changes in airplane
configurations and which tend to cause inadvertent
excursions of airplanes to large angles of attack and
sideslip are nonlinear static pitching-moment char-
acteristics at moderately low-lift coefficients, losses
in directional stability, dihedral effectiveness, and
roll damping at angles below the stall; and loss of
directional stability with increasing Mach number in
the supersonic range. The effect of inertia coupling
is also of major influence. These characteristics
which can cause inadvertent and uncontrolled mo-
tions and their recognition from wind-tunnel measure-
moments are discussed.


N-37434'

AEROELASTIC EFFECTS OF AERODYNAMIC
HEATING. Hugh L. Dryden and John E. Duberg.
(Presented to Fifth General Assembly of Advisory
Group for Aeronautical Research and Development,
Ottawa, Canada, June 10-17, 1955) 18p. diagrs.

The design of aircraft to withstand aeroelastic dif-
ficulties at high supersonic speeds will of necessity
require the consideration of the effects of aero-
dynamic heating. Among the various aeroelastic
consequences of aerodynamic heating, the reduction
of over-all stiffness through the action of thermal





NACA
RESEARCH ABSTRACTS NO. 86


stress is the most novel and may well turn out to be
the most serious. An appreciation of this phenome-
non must become part of the working equipment of
the modern aeroelastician.


N-37552*

THE SIMULATION AND MEASUREMENT OF AERO-
DYNAMIC HEATING AT SUPERSONIC AND HYPER-
SONIC MACH NUMBERS. Jackson R. Stalder and
Alvin Seiff. (Presented to Wind Tunnel and Model
Testing Panel of Advisory Group for Aeronautical
Research and Development, Ottawa, Canada,
June 10-14, 1955) 25p. diagrs., photos.

The purpose of this paper is to review the experience
of the Ames Laboratory of the NACA with respect to
the simulation and measurement of aerodynamic heat
transfer at supersonic and hypersonic speeds.


UNPUBLISHED PAPERS


N-37541*

GYROPLANES (HELICOPTERS) WITH ROTORS
THERMOPROPELLED BY LOW-PRESSURE FLUID.
(Giravions a rotors thermopropulses par fluide
moteur a basse pression). M. Rend Dorand. June
1955. 52p. diagrs., photo., tabs. (Trans. from
Technique et Science Aeronautiques, no. 3, 1951,
p. 158-189)


The propelling fluid has a compression ratio of from
1.15 to 1.4; temperature of the order of 1300 C; and
velocity of from 200 m/sec to 300 m/sec. It is pro-
duced by a generator placed in the fuselage and dis-
tributed by means of a rotating joint in the longerons
of the rotor and escapes by means of streamline
propulsion nozzles in the blade profile whose orifices
are distributed over span and depth of the blade.
Part of the propelling fluid is used for directional
maneuvering, and in certain cases, for direct pro-
pulsion of the machine. Expensive power transmis-
sion machinery is eliminated and fuel econon y re-
sults.


N-37596*

ESTIMATION OF ERRORS IN THE APPROXIMATE
SOLUTION OF LINEAR PROBLEMS. (Olsenki
pogreshnostei priblizhennykh resheni lineinykh
zadach). M. G. Slobodyanskii. 23p. (Trans. from
Prikladnaya Matematika i Mekhanika, v. 17, no. 2,
Mar. -Apr., 1953, p. 229-244)

The present paper is devoted to the further develop-
ment of the method of constructing the approximate
solution and estimation of the error in linear prob-
lems reducible to variational problems. The re-
sults are applied to the boundary problem for the
ordinary differential equation, and certain numeri-
cal examples are considered.


NACA Langley Field. Va.






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