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 )
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serial   ( sobekcm )
federal government publication   ( marcgt )
abstract or summary   ( marcgt )

Notes

Statement of Responsibility:
National Advisory Committee for Aeronautics.
Dates or Sequential Designation:
Abstracts no. 1 (June 15, 1951)-no. 93 (Nov. 30, 1955).

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 001469326
notis - AGY1019
oclc - 01471285
lccn - 86657025
issn - 0499-9274
Classification:
lcc - TL501 .U5895
System ID:
AA00009235:00040

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National Advisory Committee For Aeronautics


Research Abstracts


NO.77


FEBRUARY .2, --195S


CURRENT NACA REPORTS

NACA Rept. 1143

A VECTOR STUDY OF LINEARIZED SUPERSONIC
FLOW APPLICATIONS TO NONPLANAR
PROBLEMS. John C. Martin. 1953. i, 34p.
diagrs., tab. (NACA Rept. 1143. Formerly .
TN 2641)

A vector study of the partial-differential equali6n i~
steady linearized supersonic flow is presented.' n-
eral expressions are derived which relate thd ve.lo%-
ity potential in the stream to the conditions aonthl AR ]
disturbing surfaces. Problems concerning non-
planar systems are investigated, and methods are
derived for the solution of some simple problems.
The damping in roll is found for rolling tails coR-
sisting of four, six, and eight rectangular fins.



NACA Rept. 1152

THEORY AND PROCEDURE FOR DETERMINING
LOADS AND MOTIONS IN CHINE-IMMERSED
HYDRODYNAMIC IMPACTS OF PRISMATIC
BODIES. Emanuel Schnitzer. 1953. ii, 29p.
diagrs. (NACA Rept. 1152. Formerly TN 2813)
A theoretical method is derived for computing the
motions and hydrodynamic loads during water land-
ings of prismatic bodies involving appreciable im-
mersion of the chines. A simplified method of com-
putation covering flat-plate and V-bottom bodies with
beam-loading coefficients greater than unity is given
as a separate section. Comparisons of theory with
experiment are presented as plots of impact lift coef-
ficient and maximum draft-beam ratio against flight-
path angle and as time histories of loads and motions.
Fair agreement is found to exist for chine-immersed
landings for angles of dead rise of 00 and 300, beam-
loading coefficients from 1 to 36.5, flight-path
angles from 20 to 900, and trims from 60 to 450.


NACA Rept. 1154

ANALYSIS OF LANDING-GEAR BEHAVIOR.
Benjamin Milwitzky and Francis E. Cook. 1953.
iii, 45p. diagrs., photo., 3 tabs. (NACA Rept. 1154.
Formerly TN 2755)

The behavior of the conventional oleo-pneumatic land-
ing gear during impact is analyzed. The applicability
of the analysis to actual landing gears is established
by comparing calculated results with drop-test data.
In addition to the more exact treatment, studies are
made to determine the effects of variations in such
parameters as the force-deflection characteristics of
the tire, the orifice discharge coefficient, and the


polytropic exponent for the air-compression process
in the shock strut, which may not be known accurately
in practical design problems. An investigation is
also made to determine the extent to which represen-
tation of the system can be simpified and still yield
useful results. Generalized solutions for the behav-
ior of a simplified system, wtuch may be useful in
preliminary design, are presented for a wide range
"alanding-gear and impact parameters.

,-

NA A \ ept. 1159
1 1956
IMP' FMEMENT OF WATER DROPLETS ON WEDGES
AN-"DbUBLE-WEDGE AIRFOILS AT SUPERSONIC
,3PE)S. John S. Serafini. 1954. 11, 24p. diagrs.
(NACA Rept. 1159. Formerly TN 2971)
1


An analytical solution has been obtained for the
equations of motion of water droplets impinging on
a wedge in a two-dimensional supersonic flow field
with a shock wave attached to the wedge. The
closed-form solution yields analytical expressions
for the equation of the droplet trajectory, the local
rate of impingement and the impingement velocity
at any point on the wedge surface, and the total rate
of impingement. The analytical expressions are
utilzed to determine the impingement on the for-
ward surfaces of diamond airfoils in supersonic
flow fields with attached shock waves. The results
presented include the following conditions: droplet
diameters from 2 to 100 microns, pressure altitudes
from sea level to 30,000 feet, free-stream static
temperatures from 4200 to 4600 R. free-stream
Mach numbers from 1.1 to 2.0. semiapex angles
for the wedge from 1.140 to 7.970, thickness-to-
chord ratios for the diamond airfoil from 0.02 to
0.14, chord lengths from 1 to 20 feet, and angles
of attack from zero to the inverse tangent of the
airfoil thickness-to-chord ratio.



NACA RM A54I23

A HEATED-WIRE LIQUID-WATER-CONTENT
INSTRUMENT AND RESULTS OF INITIAL FLIGHT
TESTS IN ICING CONDITIONS. Carr B. Neel.
January 1955. 33p. diagrs., photos., Lab. (NACA
RM A54123)

A flight model of the heated-wire instrument was
tested in natural icing conditions, and was shown to
provide reliable measurements of liquid-water
content. The rapid response of the instrument
enabled detailed study of cloud structure. Cloud-
duct tests showed measurements could be made up to
700 mph. Results of the flight measurements sub-
stantiated the high values of water content previously
predicted. The highest value measured was 3.7
grams per cubic meter.


*AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 15:2 H ST, NW., WASHINGTON 25, D C, CITING CODE NUMBER ABOVE EACH TITLE;
THE REPORT TITLE AND AUTHOR.
d**tiPt 4-.






2


NACA TM 1354

GENERAL THEORY OF CONICAL FLOWS AND ITS
APPLICATION TO SUPERSONIC AERODYNAMICS.
(La th6orie ge4nrale des movements coniques et
ses applications a l'aerodynamique supersonique).
Paul Germain. PREFACE. M. J. Peres.
January 1955. vii, 333p. diagrs. (NACA
TM 1354. Trans. from Office National d'itudes et
de Recherches Adronautiques, Pub. 34, 1949)

The report deals with a method of studying the
equation of cylindrical waves particularly indicated
for the solution of certain aerodynamic problems.
The method reduces problems of a hyperbolic equa-
tion to problems of harmonic functions. The study
has been applied toward setting up the fundamental
principles, to developing their investigation up to
calculation of the pressures on the visualized ob-
stacles, and to showing how the initial field of
"conical flows" was considerably enlarged by a
procedure of integral superposition.

NACA TN 3149

PREDICTION OF LOSSES INDUCED BY ANGLES OF
ATTACK IN CASCADES OF SHARP-NOSED BLADES
FOR INCOMPRESSIBLE AND SUBSONIC COMPRES-
SIBLE FLOW. James J. Kramer and John D.
Stanitz. January 1955. 45p. diagrs. (NACA
TN 3149)

A method of computing the losses in total pressure
caused by a nonzero angle of attack at the inlet to a
row of sharp-nose4 blades is developed for both in-
compressible and subsonic compressible flow. The
results of the analysis indicate for the range of
variables considered that increases in upstream flow
angle cause sharp rises in total-pressure loss coef-
ficient and corresponding drops in static-pressure
coefficient for negative angles of attack, but for
positive angles of attack and upstream flow angles
less than 600 there is little variation in total-
pressure loss coefficient with upstream flow angle.


NACA TN 3264

STUDY OF THE MOMENTUM DISTRIBUTION OF
TURBULENT BOUNDARY LAYERS IN ADVERSE
PRESSURE GRADIENTS. Virgil A. Sandborn and
Raymond J. Slogar. January 1955. 79p. diagrs.,
photos. (NACA TN 3264)

Experimental evaluation and analysis were made of
mean and turbulent terms of the equations of motion
and the stress tensor at four stations in a turbulent
boundary layer with a progressively increasing
adverse pressure gradient. Evaluation of terms of
the stress tensor indicated that pv2, pw2, and
-puv are of equal order of magnitude, while pu2
is roughly four times larger near the wall. The

term v2 of the y-direction equation of motion
l7y
was found to be as large as any term of the x-
direction equation.

NACA TN 3266

EXPERIMENTAL EVALUATION OF MOMENTUM
TERMS IN TURBULENT PIPE FLOW. Virgil A.
Sandborn. January 1955. 40p. diagrs. (NACA
TN 3266)


NACA
RESEARCH ABSTRACTS NO. 77
Terms of the longitudinal- and radial-direction
turbulent momentum equations were experimentally
evaluated in a 4-inch-diameter pipe from total- and
static-pressure data and hot-wire anemometer
surveys. Terms of the r-direction momentum
equation were found to be as large as terms of the
x-direction equation. Direct comparisons were
made with turbulence measurements obtained using
the constant-current and constant-temperature
systems of hot-wire anemometry. The two systems
agree equally well within the experimental accuracy
of the measurements.

NACA TN 3311

DESCRIPTION AND ANALYSIS OF A ROCKET-
VEHICLE EXPERIMENT ON FLUTTER INVOLVING
WING DEFORMATION AND BODY MOTIONS. H. J.
Cunningham and R. R. Lundstrom. January 1955.
26p. diagrs., photos., 2 tabs. (NACA TN 3311.
Formerly RM L50I29)

Flight tests and a mathematical analysis were car-
ried out to demonstrate and confirm a type of sub-
sonic flutter involving rigid-body motions and wing
deformations. For the configuration considered, the
period of the oscillation was approximately 100
chords per cycle which is well within the range of
period found in dynamic-stability work on rigid air-
craft with free controls. A mathematical analysis
based on two-dimensional incompressible flow pro-
vided a conservative prediction of the airspeed at
which the low-frequency flutter occurred. Wing-
bending stiffness is the important parameter for
preventing such flutter.

NACA TN 3318

ON THE SMALL-DISTURBANCE ITERATION
METHOD FOR THE FLOW OF A COMPRESSIBLE
FLUID WITH APPLICATION TO A PARABOLIC
CYLINDER. Carl Kaplan. January 1955. 36p.
diagrs., tab. (NACA TN 3318)

The Prandtl-Busemann small-disturbance method is
applied to a parabolic cylinder and compared with
the Janzen-Rayleigh or M,2-expansion solution for
the same shape. As expected, the small-disturbance
and M.2-expansion developments are but two dif-
ferent arrangements of the actual solution. Notwith-
standing this agreement, it is concluded that the
curtailed small-disturbance solution is not suitable
for the calculation of subsonic flow past a shape
(like the parabolic cylinder) which does not possess
a control parameter such as, for example, a thick-
ness coefficient. The small-disturbance solution for
the parabolic cylinder is examined from the point of
view of thin-airfoil theory. The series development
of the fluid speed at the surface in powers of the
ratio of the radius of curvature of the vertex and the
abscissa measured from the vertex agrees with the
results of second-order thin-airfoil theory.


NACA TN 3323

CHARTS FOR ESTIMATING PERFORMANCE OF
HIGH-PERFORMANCE HELICOPTERS. Alfred
Gessow and Robert J. Tapscott. January 1955.
36p. diagrs. (NACA TN 3323) .

Theoretically derived charts are presented for use
in predicting profile-drag-thrust ratios of rotors
having hinged blades with -8 twist. The charts are
considered applicable to rotor-operating conditions





NACA
RESEARCH ABSTRACTS NO. 77

in which high tip-speed ratios or large rotor angles
of attack are encountered; however, they do not in-
clude the effects of compressibility. Limit lines
showing the conditions of onset of stall are included
in the charts, and the effects of blade twist on the
stall limits are discussed.


NACA TN 3324

A NOTE ON THE DRAG DUE TO LIFT OF REC-
TANGULAR WINGS OF LOW ASPECT RATIO.
Edward C. Polhamus. January 1955. 24p. diagrs.
(NACA TN 3324)

Methods of estimating the induced drag of low-aspect-
ratio wings are discussed and compared with experi-
ment. The profile drag due to lift is also discussed
and a method is developed which relates the effect of
aspect ratio on the profile drag due to lift to an
"effective" two-dimensional lift coefficient. A
simple expression for this effective two-dimensional
lift coefficient in terms of the aspect ratio is derived
and used to correlate experimental values of profile
drag due to lift for rectangular wings in the low-
aspect-ratio range.


NACA TN 3331

ANALYSIS OF LAMINAR FORCED-CONVECTION
HEAT TRANSFER IN ENTRANCE REGION OF FLAT
RECTANGULAR DUCTS. E. M. Sparrow. January
1955. 42p. diagrs. (NACA TN 3331)

The simultaneous development of temperature and
velocity profiles in the entrance region of a flat rec-
tangular duct is studied. The flow is laminar with
constant properties and negligible dissipation. Two
thermal conditions for the duct walls are considered:
(1) both walls have the same uniform temperature
throughout; (2) one wall is at uniform temperature,
the other wall is insulated. Thermal and velocity
boundary layers are calculated using the Karman-
Pohlhausen method. Nusselt numbers are reported
for Prandtl number in the range 0.01 to 50. For
simultaneously developing profiles, the Nusselt num-
ber is found not to be a function at Graetz number
alone, as it is for an unchanging parabolic velocity
profile throughout.


NACA TN 3335

METHODS FOR RAPID GRAPHICAL EVALUATION
OF COOLED OR UNCOOLED TURBOJET AND
TURBOPROP ENGINE OR COMPONENT PERFORM-
ANCE (EFFECTS OF VARIABLE SPECIFIC HEAT
INCLUDED). Jack B. Esgar and Robert R. Ziemer.
January 1955. 45p. diagrs. (NACA TN 3335)

Curves based on the thermodynamic properties of air
and combustion gases for a hydrogen-carbon ratio of
0.167 are presented to relate parameters affecting
each engine component. The curves cover a rangeof
flight Mach numbers from to 3.0, compressor pres-
sure ratios from 1 to 30, turbine-inlet temperatures
from 15000 to 30000 R, and afterburner temperatures
from 28000 to 35000 R. Except for extreme cases,
the curves are accurate to at least 30 R in tempera-
ture and I percent in pressure ratio, fuel-air ratio,
and specific thrust. Procedures required for per-
formance evaluation are explained for both uncooled


3


engines with no compressor bleed and for engines
utilizing both compressor bleed and turbine cooling.


NACA TN 3337

INVESTIGATION OF TEMPERATURE LIMITATION
OF VARIOUS LUBRICANTS FOR HIGH-
TEMPERATURE 20-MILLIMETER-BORE BALL
BEARINGS. Z. N. Nemeth and W. J. Anderson.
January 1955. 31p. diagrs., photos., 2 tabs.
(NACA TN 3337)

Twenty-millimeter-bore tool-steel ball bearings,
equipped with either a beryllium copper or an Inconel
cage, were operated with liquid and with solid
lubricants at temperatures from 1000 to 10000 F at a
speed of 2500 rpm and a thrust load of 110 lb. Solid
lubricants were more effective than fluid lubricants
at the higher temperatures. Graphite provided ef-
fective lubrication to 10000 F with bearings equipped
with either a beryllium copper or an Inconel cage;
molybdenum disulfide, to 8500 F with a bearing
equipped with an Inconel cage. A silicone-diester
blend, the best high-temperature liquid lubricant,
provided effective lubrication to 7000 F and allowed
operation of the bearing at 8500 F although the bear-
ing operation was rough and friction torque was high.


NACA TN 3346

PREDICTION OF DOWNWASH BEHIND SWEPT-WING
AIRPLANES AT SUBSONIC SPEED. John DeYoung
and Walter H. Barling, Jr. January 1955. 104p.
diagrs., 3 tabs. (NACA TN 3346)

The numerical integration method presented enables
a rapid prediction of downwash. The principal effects
of the rolling-up of the wake are treated as correc-
tions to the flat-sheet wake. A simple approximate
correction for the effect of the fuselage is applied.
Computing forms and charts of pertinent functions
are included. Agreement with available experimen-
tal data is good.


NACA TN 3353

EFFECTIVE MOMENT OF INERTIA OF FLUID IN
OFFSET, INCLINED, AND SWEPT-WING TANKS
UNDERGOING PITCHING OSCILLATIONS. James
R. Reese and John L. Sewall. January 1955. 27p.
diagrs., 6 tabs. (NACA TN 3353)
Fluid-dynamics studies were made of simplified
model fuel tanks undergoing pitching oscillations.
The tanks were pylon mounted, centrally mounted at
angles of sweep, and inclined at angles of attack.
The effective moment of inertia of the fluid was
determined experimentally for the various tank con-
figurations over a tank-fullness range from empty to
full. For full pylon-mounted and swept-wing tanks,
comparisons of experimental and theoretical solu-
tions for the effective moment of inertia of fluid
showed good agreement. Studies of the effect of
vertical, horizontal, and diffused baffles in pylon-
mounted tanks showed that the effective moment of
inertia of the fluid was considerably less without
baffles. Diffused baffles were found to have high
damping characteristics in pylon-mounted tanks and
very low damping characteristics in centrally
mounted tanks.






4


NACA TN 3356

EFFECT OF LAG OF SIDEWASH ON THE
VERTICAL-TAIL CONTRIBUTION TO OSCIL-
LATORY DAMPING IN YAW OF AIRPLANE
MODELS. Lewis R. Fisher and Herman S.
Fletcher. January 1955. 38p. diagrs., photos.
(NACA TN 3356)

Two models were tested which permitted, in effect,
a systematic variation of the sidewash gradient at
the vertical tail. For the first model, this variation
was accomplished by mounting auxiliary fins forward
of the vertical tail; for the second model, it was
done by altering the vertical location of the wing on
the fuselage. The unsteady damping-in-yaw and
directional-stability parameters are compared with
the steady derivatives obtained for the same models
to establish the effects of the sidewash and the lag
of the sidewash on these lateral stability derivatives.

NACA TN 3357

THE EFFECTS OF VARIOUS PARAMETERS,
INCLUDING MACH NUMBER, ON PROPELLER-
BLADE FLUTTER WITH EMPHASIS ON STALL
FLUTTER. John E. Baker. January 1955. 40p.
diagrs., 3 tabs. (NACA TN 3357. Formerly
RM L50L12b)
The effect of many parameters significant to wing
flutter as well as blade twist was studied on several
untwisted rotating models to determine their signif-
icance with respect to propeller stall flutter. The
minimum values of the flutter-speed coefficient were
found to be slightly greater than 1.0 at subcritical
Mach numbers. Of the few parameters that raised
the minimum flutter-speed coefficients, forward
movement of the section center-of-gravity location
and Mach number at supercritical speeds were most
significant. The effect of Mach number was of such
significance that a tentative criterion for designing
completely flutter-free thin supersonic propellers is
indicated.


NACA TN 3360

SOME EFFECTS OF PROPELLER OPERATION AND
LOCATION ON ABILITY OF A WING WITH PLAIN
FLAPS TO DEFLECT PROPELLER SLIPSTREAMS
DOWNWARD FOR VERTICAL TAKE-OFF. John W.
Draper and Richard E. Kuhn. January 1955. 28p.
diagrs., photo. (NACA TN 3360)

An investigation has been conducted of the effects of
propeller blade angle, mode of propeller rotation,
propeller location, and ratio of wing chord to pro-
peller diameter on the ability of a wing with plain
flaps to deflect the propeller slipstream downward in
order to achieve vertical take-off. The basic model
consisted of a semispan wing with 30-percent-chord
and 60-percent-chord plain flaps. Two large-
diameter overlapping propellers driven by electric
motors were used.


NACA TN 3361

AERODYNAMIC CHARACTERISTICS OF NACA 0012
AIRFOIL SECTION AT ANGLES OF ATTACK FROM
00 TO 1800. Chris C. Critzos, Harry H. Heyson
and Robert W. Boswinkle, Jr. January 1955. 21p.
diagrs. (NACA TN 3361)


NACA
RESEARCH ABSTRACTS NO.77

The aerodynamic characteristics of the NACA 0012
airfoil section are presented for an angle-of-attack
range extending through 1800. Data were obtained
at a Reynolds number of 1.8 x 106 with the airfoil
surfaces smooth and with roughness applied at the
leading and trailing edges and at a Reynolds number
of 0.5 x 106 with the airfoil surfaces smooth. The
tests were conducted in the Langley low-turbulence
pressure tunnel at Mach numbers no greater than
0.15.


NACA TN 3364

INVESTIGATION OF EFFECTIVENESS OF LARGE-
CHORD SLOTTED FLAPS IN DEFLECTING
PROPELLER SLIPSTREAMS DOWNWARD FOR
VERTICAL TAKE-OFF AND LOW-SPEED FLIGHT.
Richard E. Kuhn and John W. Draper. January
1955. 42p. diagrs., photo., tab. (NACATN 3364)

An investigation of the effectiveness of a wing
equipped with large-chord slotted flaps and an aux-
iliary vane in rotating the effective thrust vector of
propellers to a near-vertical direction for vertical
take-off and low-speed flight has been conducted.
The model consisted of a semispan wing equipped
with 60-percent-chord and 30-percent-chord slotted
flaps. Two large-diameter overlapping propellers,
driven by electric motors, were used. The effect of
wing incidence, propeller blade angle, and an aux-
iliary vane on the ability of the wing equipped with
slotted flaps to deflect the propeller slipstreams
downward were also investigated. A few tests
covered an angle-of-attack range from 00 to 900
and a thrust-coefficient range representing free-
flight velocities from zero to the normal range of
cruising velocity.




NACA TN 3366

A METHOD FOR STUDYING THE TRANSIENT
BLADE-FLAPPING BEHAVIOR OF LIFTING ROTORS
AT EXTREME OPERATING CONDITIONS. Alfred
Gessow and Almer D. Crim. January 1955. 27p.
diagrs. (NACA TN 3366)

A method is presented for studying the transient
behavior of the flapping motion, as well as for cal-
culating the steady-state flapping amplitudes, of
free-to-cone and seesaw rotors operating at extreme
flight conditions. The method is general and can he
applied to blades of any airfoil section, mass dis-
tribution, twist, plan-form taper, root cutout, and
flapping-hinge geometry. Stall and compressibility
effects can also be accounted for. Applications of
the method to the calculation of the stability of the
flapping motion of unloaded rotors and to the tran-
sient blade motion resulting from arbitrary control
inputs under conditions of extreme stall are included.


NACA TN 3387

USE OF NONLINEARITIES TO COMPENSATE FOR
THE EFFECTS OF A RATE-LIMITED SERVO ON
THE RESPONSE OF AN AUTOMATICALLY CON-
TROLLED AIRCRAFT. Stanley F. Schmidt and
William C. Triplett. January 1955. 27p. diagrs.,
tab. (NACA TN 3387)






NACA
RESEARCH ABSTRACTS NO. 77

A method is developed for designing suitable non-
linear functions of error into a system to compensate
for the undesirable effects of control-surface rate
limiting on the response of an automatically con-
Srolled aircraft.


BRITISH REPORTS


N-34602"

Aeronautical Research Council (Gt. Brit.)
THE EFFICIENCY OF HIGH-SPEED WIND TUNNELS
OF THE INDUCTION TYPE. A. E. Knowler and
D. W. Holder. APPENDIX THE EFFICIENCY OF
INTERMITTENT OPERATION FROM COMPRESSED
AIR STORAGE. D. W. Holder. 1954. iii, 59p.
diagrs., photo., 6 tabs. (ARC R & M 2448.
Formerly ARC 7563; 7672; 7756; 7812; 8024;
8138; 8394; 8668; 8669; 8670; 9490 and 9902)

An investigation has been made of the influence of
various design features upon the efficiency of an
induction-type tunnel with an annular injector slot. A
model tunnel of circular cross-section was used, the
working-section diameter being 2-1/4 inch. At sub-
sonic working-section speeds it was found that effi-
ciencies comparable with those of direct-action tun-
nels could be attained, but at supersonic speeds the
efficiency began to fall rapidly. A comparison be-
tween the 2-1/4 inch and 12 inch tunnels is made.


N-34603*

Aeronautical Research Council (Gt. Brit.)
ASSESSMENT OF THE RELATIVE PERFORMANCE
OF THE BY-PASS ENGINE AND THE ORTHODOX
DOUBLE COMPOUND JET ENGINE. E. A. Bridle.
1954. 12p. diagrs. (ARC R & M 2862; ARC 11,740.
Formerly NGTE Memo. M. 32)

The by-pass engine can be described as a form of
ducted fan engine in which the fan boosts the main
compressor. Two possible forms of by-pass engine
are described, and their estimated performance is
compared with that of the orthodox double compound
jet engine under various flight conditions, the cal-
culations being extended to include the case of thrust
boosting by means of exhaust reheat. It is concluded
that the by-pass engine can offer an appreciable gain
in respect of fuel economy over the orthodox double
compound jet engine even at 650 mph in the strato-
sphere, at the expense, however, of increased
frontal area for a given thrust.

N-34604*

Aeronautical Research Council (Gt. Brit.)
FULL-SCALE MEASUREMENTS OF IMPACT LOADS
ON A LARGE FLYINGBOAT. PART 1. DESCRIPTION
OF APPARATUS AND INSTRUMENT INSTALLATION-
J. W. McIvor. 1954. 27p. photos., diagrs.
(ARC CP 182)

The variations with time of the total force and the
distribution of water pressures on the hull bottom of
a flying boat are related to the horizontal velocity,
vertical velocity and keel attitude relative to the water
during impact. Methods are described for obtaining,
in a form suitable for easy analysis of results,
records of these variables, in order to verify impact


theories. The equipment used included transducers
for the conversion of physical quantities to electrical
signals, multichannel electronic amplifiers, and
mirror galvanometer recorders.


N-34605*

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF INLET CONDITIONS ON THE
FLOW IN ANNULAR DIFFUSERS. I. H. Johnston.
1954. 15p. diagrs. (ARC CP 178)

Tests have been carried out on annular diffusers
having a common area ratio of 3. 19 and varying in
divergence angle from 6.50 to 150. The performance
of each diffuser has been measured for a variety of
inlet velocity distributions and the effect of axially
splitting the flow in the diffusers has been investi-
gated. Diffuser efficiency is found to deteriorate as
inlet conditions become nonuniform, this tendency
increasing with diffuser angle. Splitting of the higher
angle diffusers improves efficiency for nonuniform
profiles, but these increases in efficiency are
accompanied by pronounced static pressure gradients
across the diffuser throat which in certain applica-
tions might prove undesirable.



N-34606*

Aeronautical Research Council (Gt. Brit.)
A METHOD OF COMPUTING SUBSONIC AND TRAN-
SONIC PLANE FLOWS. C. S. Sinnott. 1954. 21p.
diagrs., photos. (ARC CP 173)

A relaxation treatment of a simple but exact differen-
tial equation for compressible flow is presented. The
method has advantages over other numerical treat-
ments of the same problem and because of the sim-
plicity of the basic differential equation should be
particularly suitable for high-speed computing
machines. The flow about a 10-percent-thick airfoil
(RAE 104) at zero incidence is calculated for Mach
numbers to 0.70, 0.79, and 0.86. At M = 0.86 the
existence, but not the position of a transonic shock
wave is predicted by the relaxation technique. Satis-
factory agreement with experiment is obtained.


N-34607*

Aeronautical Research Council (Gt. Brit.)
THE THEORETICAL INTERFERENCE VELOCITY ON
THE AXIS OF A TWO-DIMENSIONAL WIND TUNNEL
WITH SLOTTED WALLS. R. C. Tomlinson. 1954.
15p. diagrs., 2 tabs. (ARC CP 181)

A calculation is made of the interference velocity on
the axis in the flow of an incompressible fluid past a
line doublet in a two-dimensional slotted-wall tunnel,
i. e. a rectangular tunnel whose shorter sides are
slotted in the direction on the flow. Numerical solu-
tions are obtained which show that if the width of
(slot + slat) is one twelfth of the tunnel height (e. g.
six slots in the shorter side of a 2 by 1 tunnel), the
interference velocity is little different from the cor-
responding open-jet value i. e. no slats for all
slot/slat ratios greater than 1/40. It appears likely
that all cases which give conditions uniform across
the center plane of the tunnel will also give an inter-
ference velocity which is close to the open jet figure.





6



N-34608*

Aeronautical Research Council (Gt. Brit.)
THE M. A. E. E. RECORDING ACCELEROMETER.
D. M. Ridland and R. Parker. 1954. 20p. photos.,
diagrs. (ARC CP 177)

The M. A. E. E. recording accelerometer is basically
the accelerometer unit of a desynn accelerometer,
adapted to make a continuous and immediate presen-
tation of accurate, calibrated accelerations on a half
second time base. The recording medium is metal-
lized paper, having a speed of half an inch per sec-
ond, and the instrument can be operated continuously
for 20 minutes on one loading. It can record with
full scale deflections, from 1 to 10gg when the natu-
ral frequencies will be about 7 and 22 cps, respec-
tively. The instrument is simple, it has been proved
reliable and accurate and is most convenient in use.

N-34616*

Nat. Gas Turbine Establishment (Gt. Brit.)
THERMODYNAMIC PROPERTIES OF AIR AND
COMBUSTION PRODUCTS OF HYDROCARBON
FUELS. PART II. THE FLOW OF GASES OF
VARYING SPECIFIC HEAT. D. Fielding, J. E. C.
Topps and W. R. Thomson. June 1954. 43p. diagrs.
(NGTE R. 160)

General methods of calculating gas flow have been
investigated and it has been found possible to present
data in graphical form which permits the frictionless
flow properties of ideal gases of varying specific
heat and any molecular weight to be calculated at
velocities up to sonic with error not exceeding that
inherent in the approximation to four significant
figures. Charts are presented of a total head flow
parameter against pressure ratio, pressure ratio
against temperature ratio, and Mach number against
temperature ratio, all with molar specific heat as a
running parameter. The relations are also given for
Mach numbers 1.0 to 3.0.

N-34617 *
Nat. Gas Turbine Establishment (Gt. Brit.)
TESTS TO DETERMINE THE EFFECT OF THE EX-
HAUST CONE SUPPORT STRUT FAIRINGS ON THE
PERFORMANCE OFA TURBO-JET ENGINE. W.
Deacon. July 1954. 19p. diagrs. (NGTE Memo. M. 222)

Simple bench tests have been made to determine what
effect the exhaust cone support strut fairings have on
the performance of a turbojet engine. Pitot and yaw-
meter traverses were made just downstream of the
turbine. The results show that some residual swirl
is present in the turbine exhaust, and that the fairings
effectively remove it. Removal of the strut fairings
causes an increase in exhaust losses, mainly due to
the increased swirl in the exhaust system, but also
due to the exposure of unfaired support rods. The
increased loss due to swirl is in accordance with that
predicted from model tests. When the fairings are
removed, removal of the bullet causes no further
increase in exhaust losses.


N-34619*

Nat. Gas Turbine Establishment (Gt. Brit.)
SOME CALCULATIONS ON IDEAL COMBUSTION IN
A PARALLEL DUCT. A. B. P. Beeton. July 1954.
25p. diagrs. (NGTE Memo.M. 221)


NACA
RESEARCH ABSTRACTS NO. 77


Taking one typical case representative of ram-jet
combustion conditions, calculations were made of
the integrated kinetic energy and throat specific
impulse after loss-free combustion in a parallel
duct. Various inlet flow velocities were considered,
corresponding to downstream throat areas between
83-1/2 percent and 95-1/2 percent of the duct area.
Comparing these results with those calculated for
the same approach velocities on the usual basis of
homogeneous flow, the latter are shown to involve a
loss in specific impulse of at most half a second in
150. The corresponding kinetic energies were not
calculated with sufficient accuracy to obtain more
than the order of the loss, which appeared to be
about half the energy in the approach flow.

N-34621*

Nat. Gas Turbine Establishment (Gt. Brit.)
THE EFFECT OF BURNER SHROUD AIR ON A
FUEL SPRAY. H. Clare. July 1954. 25p. diagrs.,
photos. (NGTE Memo. M. 223)
The possibility of using shroud air as an atomization
booster has been investigated. It was found to be ef-
fective only when the initial spray was coarse. Drop
size analyses of sprays produced from swirl atom-
izers at low fuel pressures showed that at low shroud
air pressure swirling air was the more effective.
At higher air pressures of 20 in. water the effects of
both swirled and unswirled air were similar. The
difference between the swirled and unswirled air was
most marked at a pressure of 2 in. head of water
where the swirling air improved atomization, whilst
unswirled air produced a still coarser spray. It is
concluded that swirling shroud air could be used to
advantage where both fuel and shroud air pressures
were low, a condition which might be expected in an
aircraft combustion chamber idling at altitude.


N-34623*

Nat. Gas Turbine Establishment (Gt. Brit.)
A CORRECTED SPEED TACHOSCOPE. R. Staniforth.
May 1954. 15p. diagrs., photos. (NGTE Memo. M.217)
It is often desirable in testing aerodynamic compres-
sors to take all readings at fixed corrected speeds
rather than true speeds. This corrected speed is
defined as the actual shaft speed divided by the
square root of the ratio of the absolute air inlet
temperature to the standard temperature (2880 K.).
With the instrument described, this is possible with
high accuracy (error < 0.1 percent) without further
complication than the setting of a dial to the observed
air inlet temperature. The latter operation could be
dispensed with and the correction obtained directly
from a temperature sensitive device such as a
thermistor or a resistance thermometer element.

N-34627*

Marine Aircraft Experimental Establishment
(Gt. Brit.) SOME ASPECTS OF THE FLOW ROUND
PLANING SEAPLANE HULLS OR FLOATS AND
IMPROVEMENT IN STEP AND AFTERBODY
DESIGN. K. M. Tomaszewski and A. G. Smith.
September 1954. 13p. diagrs., photo. (MAEE
F/TN/4; ARC 14,376; ARC S. 669. Formerly MAEE
Tech. Memo 5)

A method of step and afterbody design is given and
illustrated which is based on knowing the shape of
the wake produced by the forebody and the interaction






NACA
RESEARCH ABSTRACTS NO. 77


between air and water flow under the afterbody. It
is claimed that this is a logical approach which gives
the designer a clear overall picture of the physical
conditions existing during take-off and landing,
enabling Ihe optimum aero- and hydrodynamic per-
formance to be quickly obtained. Application leads
to large improvements in air and water drag and in
water stability.



N-34628*

Royal Aircraft Establishment (Gt. Brit.)
CALIBRATION OF THE FLOW IN THE WORKING
SECTION OF THE 3 FT. x 3 FT. TUNNEL,
NATIONAL AERONAUTICAL ESTABLISHMENT.
D. E. Morris. September 1954. 44p. diagrs.
(RAE Tech. Note Aero 2336)

A number of calibrations, consisting of both pitot
and static pressure measurements and also flow
direction measurements, have been made of the flow
in the working section of the 3-ft by 3-ft supersonic
tunnel. In this report some selected examples are
given to show the general nature of the flow distri-
bution with the M = 1.4, 1.6, 1.8, and 2.0 nozzles
and to demonstrate a number of interesting points
in the measurements and in the characteristics of
the flow.





N-34630*

Ministry of Supply (Gt. Brit.)
REPORT ON PROGRESS IN THE DEVELOPMENT
OF AN ALLOYING THEORY OF TITANIUM. Dept.
of Physical Metallurgy, Birmingham University.
(Progress report 1951-1954 of the titanium research
programme) November 1954. 28p. diagrs., photos.
(MOS S & TM 13/54)

In this report an attempt has been made to summa-
rize those properties of titanium and its alloys which
would appear to have some significant relationship to
the electronic structure of the metal and to put for-
ward tentative suggestions for possible causes of the
observed behavior of its alloys.





N-34631*

M inst ry of Supply (Gt. Brit.)
DIFFUSION OF SOLUTE ELEMENTS IN TITANIUM.
D. H. Tomlin and A. J. Mortlock. (Biannual
progress report no. 1 for April-October 1954).
November 1954. 6p. diagrs. (MOS S & TM 14/54)

The initial part of a general study of the diffusion of
solule elements in titanium reported here has been
aimed at experimentally determining the diffusion
parameters associated with the movements of chro-
mium in 0-titanium. A new method is described
which makes use of autoradiographic techniques, and
which suggests that a number of diffusion coefficients
corresponding to different temperatures may be
determined using the same specimen.


7



MISCELLANEOUS


NACA Rept. 1152

Errata on "THEORY AND PROCEDURE FOR
DETERMINING LOADS AND MOTIONS IN CHINE-
IMMERSED HYDRODYNAMIC IMPACTS OF
PRISMATIC BODIES. Emanuel Schnitzer. 1953.

NACA Rept. 1159

Errata on "IMPINGEMENT OF WATER DROPLETS
IN WEDGES AND DOUBLE-WEDGE AIRFOILS AT
SUPERSONIC SPEEDS. John S. Serafini. 1954.


N-33959*

Advisory Group for Aeronautical Research and
Development. A SUMMARY OF THE TECHNIQUES
OF VARIABLE MACH NUMBER SUPERSONIC
WIND TUNNEL NOZZLE DESIGN. J. T. Kennedy
and L. M. Webb. October 1954. 133p. diagrs.,
photos. (AGARDograph 3)

This report is a survey of the techniques of two-
dimensional wind-tunnel nozzle design. A procedure
for the aerodynamic design of flexible nozzles capa-
ble of continuous Mach number variation is devel-
oped in detail. The special structural, mechanical,
calibration, and cost estimation problems involved
in flexible nozzle construction are discussed.



N-35025*

Advisory Group for Aeronautical Research and
Development. METHODS AND CRITERIA FOR THE
SELECTION OF FLYING PERSONNEL. (Symposium
held February 23-25, 1953, Paris) December 1954.
59p. (AGARDograph 2)

Original papers and summaries of the symposium
covering three different aspects of the problem:
(1) psychological and psychiatric methods of selec-
tion and assessment of flying personnel; (2) clinical
selection criteria and methods; (3) physiological
selection criteria and methods are presented.


DECLASSIFIED NACA REPORTS


NACA RM L7L12

MEASUREMENTS OF THE WING AND TAIL LOADS
DURING THE ACCEPTANCE TESTS OF BELL XS-1
RESEARCH AIRPLANE. De E. Beeler and John P.
Mayer. April 13, 1948. 25p. diagrs., photos., 2
tabs. (NACA RM L7L12) (Declassified from
Confidential, 12/10/54)

During the acceptance tests of the XS-1 airplane,
strain-gage measurements were made of wing and
tail loads up to a Mach number of 0.80. The maxi-
mum lift and buffet boundaries were also determined.
The loads encountered were well within the design
loads and showed fairly good agreement with wind-
tunnel and calculated data.




UNIVERSITY OF FLORIDA

I II62 l815111
3 1262 08153 111 2


NACA RM L8A09

GENERAL HANDLING-QUALITIES RESULTS
OBTAINED DURING ACCEPTANCE FLIGHT TESTS
OF THE BELL XS-1 AIRPLANE. Walter C.
Williams, Charles M. Forsyth and Beverly P.
Brown. April 19, 1948. 72p. diagrs., photos., tab.
(NACA RM L8A09) (Declassified from Confidential,
12/10/54)
Presents results obtained during acceptance tests of
the Bell XS-1 airplane from minimum speed to a
Mach number of 0.8. The longitudinal stability was
positive but low for all conditions tested. The
directional stability was exceptionally high, and the
lateral stability was positive throughout the speed
range tested. Aileron control was adequate. The
stalling characteristics were considered satisfactory
with stall warning in the form of buffeting. The
large amount of friction in the control system,
coupled with low control-operating forces, made
exact control of the airplane difficult.


NACA RM L50J27

STABILITY AND CONTROL CHARACTERISTICS OF
A 1/4-SCALE BELL X-5 AIRPLANE MODEL IN
THE LANDING CONFIGURATION. Robert E. Becht.
December 18, 1950. 38p. diagrs., photos. (NACA
RM L50J27) (Declassified from Confidential,
10/29/54)

An investigation was conducted to determine the
stability and control characteristics of a 1/4-scale
model of a preliminary Bell X-5 airplane design in
the landing configuration with and without dive
brakes. Tests were made at wing sweep angles of
200 and 600.


NACA RM L50K30

THE ORIGIN OF AERODYNAMIC INSTABILITY OF
SUPERSONIC INLETS AT SUBCRITICAL CONDI-
TIONS. Antonio Ferri and Louis M. Nucci.
January 26, 1951. 1llp. diagrs., photos., tab.
(NACA RM L50K30) (Declassified from
Confidential, 1/11/55)


NACA
RESEARCH ABSTRACTS NO. 77


The starting of "buzz' is explained as being caused
by a separation on the inner surface of the cowling
produced by a velocity discontinuity across a vortex
sheet arising from a shock intersection. The
explanation has been confirmed by tests of a number
of inlet designs, and it has been shown that a useful
range of stable flow regulation is attainable. Values
of minimum stable entering volume flow for configur-
ations of practical interest are presented for a range
of Mach number from 1.90 to 2.70.



NACA RM L53E06a

SOME NOTES ON THE AERODYNAMIC LOADS
ASSOCIATED WITH EXTERNAL-STORE
INSTALLATIONS. H. Norman Silvers and Thomas
C. O'Bryan. June 1953. 17p. diagrs. (NACA
RM L53E06a) (Declassified from Confidential,
12/10/54)

The results of recent investigations of the aero-
dynamic loads associated with external stores are
considered with a view toward indicating certain
observations that may prove helpful in the design of
future external store installations. Consideration
was given to store-induced wing loads and to direct
external store loads.




NACA RM L53G07

WIND-TUNNEL INVESTIGATION OF THE BE-
HAVIOR OF PARACHUTES IN CLOSE PROXIMITY
TO ONE ANOTHER. Stanley H. Scher. August
1953. 12p. photos. (NACA RM L53G07)
(Declassified from Confidential, 1/11/55)

An investigation made with small tethered para-
chutes in the Langley 20-foot free-spinning tunnel
indicated that parachutes side by side had little
effect on one another, but that when one parachute
got into the wake above another parachute, the
upper parachute sometimes collapsed on top of the
lower parachute.


NACA-LanglOe 12--55 4




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