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
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Washington, D.C
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completely irregular

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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:
AA00009235:00053

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National Advisory Committee for Aeronautics


Research Abstracts


NO.89


SEPTEMBER 8, 1955


CURRENT NACA REPORTS

NACA Rept. 1190

AXIAL-LOAD FATIGUE PROPERTIES OF 24S-T
AND 75S-T ALUMINUM ALLOY AS DETERMINED
IN SEVERAL LABORATORIES. H. J. Grover and
W. S. Hyler, Battelle Memorial Institute, Paulhuhn
and Charles B. Landers, Langley Aeronautical
Laboratory and F. M. Howell, Aluminum Company
of America. 1954. ii, 25p. diagrs.. photos..
7 tabs. (NACA Rept. 1190. Formerly TN 29281
In the initial phase of an NACA program on fatigue
research, axial-load tests on 24S-T3 and 75S-T6
aluminum-alloy sheet have been made at the Battelle
Memorial Institute and at the Langley Aeronautical
Laboratory of the NACA. The test specimens were
polished and unnotched. The manufacturer of the
material, the Aluminum Company of America, has
made axial-load tests on 24S-T4 and 75S-T6 rod
material. The test techniques used at the three lab-
oratories are described in detail; the test results are
compared with each other and with results obtained
on unpolished sheet by the National Bureau of -
Standards. /



NACA RM E55F09 Ir

PRELIMINARY PERFORMANCE DATA OF RAL
'TAIL-PIPE-CASCADE-TYPE MODEL THRU \
eEVERSERS. James G. Henzel, Jr. and Jac .
McArdle. August 1955. 48p. diagrs., photos., t.
(NACA RM E55F09)

The performance of several tail-pipe-cascade-type-
model thrust reversers was obtained with sym-
metrical and asymmetrical cascade blade shapes.
One reverser with a symmetrical cascade-blade
shape had a maximum reversed thrust ratio of 76
percent of full forward thrust. This reverser also
had good thrust modulation characteristics. Another
'Vrerser but with an asymmetrical cascade blade
shape had a maximum reversed thrust ratio of 90
percent of full forward thrust.


NACA TN 3467

EFFECT OF INTERACTION ON LANDING-GEAR
BEHAVIOR AND DYNAMIC LOADS IN A FLEXIBLE
AIRPLANE STRUCTURE. Francis E. Cook and
Benjamin Milwitzky. August 1955. 75p. diagrs.,
2 tabs. (NACA TN 3467)

The effects of interaction between a landing gear and
: a flexible airplane structure on the behavior of the
landing gear and the loads in the structure have been


studied by treating the equations of motion of the
airplane and the landing gear as a coupled system.
where the landing gear is considered to have non-
linear characteristics. Numerical examples, based
on the characteristics of two large airplanes, show
the effects of interaction on the loads in the landing
gear and in the structure. The effects of neglecting
interaction and using the I;ni-ga'iijl 'unc
tion for a rigid body in a d nam- analysis oi a a"3 -
flexible airframe are also iscused. ...-

i ."
NACA TN 3474
--- -- ***
RAPID RADIANT-HEATIN TESTOppFILMLVJ [QlR
BEAMS. Joseph N. Kolant'., ..l.. ... k 4-8
Jr., and Robert D. Ross.September 1955.30p.
diagrs., photos., tab. (NACA TN 3474)

Results are described of rapid-heating tests on lour
unloaded multiweb box beams. Temperature distri-
butions and strains measured during a heating cycle
are given. An analysis of the thermal stress
required to cause buckling is in agreement with the
,. experimentally observed results. A transient heat-
",f.ng apparatus used in the tests and capable of pro-
i''Lle.Ing heating rates up to 101) Btu ft2-sec is des-
rinhed.


N TN 3478

N ATTALL BODIES OF REVOLUTION HAVING
M UMUM WAVE DRAG. Keith C. Harder and
rad Rennemann, Jr. August 1955. 28p. diagrs.,
,ab. (NACA TN 3478)

The problem of determining the shape of slender
boattail bodies of revolution for minimum wave drag
has been reexamined. It was found that minimum
solutions for Ward's slender-body drag equation can
exist only for the restricted class of bodies for which
the rate of change of cross-sectional area at the base
is zero. In order to eliminate this restriction, cer-
tain higher order terms must be retained in the drag
equation and isoperimetric relations. The minimum
problem for the isoperimetric conditions of given
length, volume, and base area is treated as an
example. According to Ward's drag equation, the
resulting body shapes have slightly less drag than
those determined by previous investigators.


NACA TN 3480

FREE-SPINNING-TUNNEL INVESTIGATION OF
GYROSCOPIC EFFECTS OF JET-ENGINE ROTAT-
ING PARTS (OR OF ROTATING PROPELLERS) ON
SPIN AND SPIN RECOVERY. James S. Bowman, Jr.
August 1955. 21p. diagrs., photo.. 2 tabs., 6 charts.
(NACA TN 3480)


C) sCVlt


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










A rotating flywheel powered by a model airplane
engine mounted on a model in the Langley 20-foot
free-spinning tunnel was used to simulate the gyro-
scopic effects of jet-engine rotating parts (or of
rotating propellers) on the erect spin and spin-
recovery characteristics of a military attack air-
plane. The effects of flywheel rotation on the angle
of attack, the rate of rotation in spins, and the
number of turns required for recovery are presented
for various control settings and for a range of
loadings.



NACA TN 3483

AN ANALYSIS OF ACCELERATION, AIRSPEED,
AND GUST-VELOCITY DATA FROM A FOUR-
ENGINE TRANSPORT AIRPLANE IN OPERATIONS
ON AN EASTERN UNITED STATES ROUTE.
Thomas L. Coleman and Mary W. Fetner. September
1955. 20p. diagrs., 3 tabs. (NACA TN 3483)

Time-history data obtained by the NACA VGH re-
corder from one model of a four-engine civil trans-
port airplane during operations on an eastern United
States route are analysed to determine the magnitude
and frequency of occurrence of gusts, gust accelera-
tions, and the associated airspeeds. Comparison of
the results with previously reported results indicates
the gust-load history to be more severe than for two
other similar operations involving other types of
four-engine transports. The present airplane was
operated in rough air at a higher percentage of its
design speed than were the other airplanes and this
condition was primarily responsible for the more
severe load history for the operation.



NACA TN 3489

CONTRIBUTIONS ON THE MECHANICS OF
BOUNDARY-LAYER TRANSITION. G. B. Schubauer
and P. S. Klebanoff, National Bureau of Standards.
September 1955. 31p. diagrs. (NACA TN 3489)

The manner in which flow in a boundary layer be-
comes turbulent was investigated on a flat plate at
wind speeds generally below 100 feet per second.
Hot-wire techniques were used, and many of the re-
sults are derived from oscillograms of velocity
fluctuations in the transition region. The more
familiar aspects of transition are discussed, and
the facts discovered while studying the character-
istics of artificially produced turbulent spots are
presented.




NACA TN 3505

AN EXPERIMENTAL INVESTIGATION OF REGIONS
OF SEPARATED LAMINAR FLOW. Donald E.
Gault. September 1955. 65p. diagrs., photos., 4 tabs.
(NACA TN 3505)

Results are given for an experimental investigation
of regions of separated flow initiated by separation
of the laminar boundary layer laminarr separation
'bubbles"). Total- and static-pressure surveys,
hot-wire anemometer observations, and detailed


NACA
RESEARCH ABSTRACTS NO. 89


pressure-distribution measurements were made in
regions of separated flow on two airfoils. The
measurements were obtained for a wide range of
angles of attack and for Reynolds numbers from 1.5
to 10 million. Some results on the effects of an in-
crease in turbulence level of the free stream are
also included.



NACA TN 3506

CRITERIONS FOR PREDICTION AND CONTROL OF
RAM-JET FLOW PULSATIONS. William H.
Sterbentz and John C. Evvard. August 1955. 60p.
diagrs., photos. (NACA TN 3506. Formerly
RM E51C27)

Results of a theoretical and experimental study of
ram-jet diffuser flow pulsing, commonly referred to
as a "buzz condition," with and without combustion
are presented. The theoretical approach to the
problem is a simplified treatment of the ram jet
likened to act as a Helmholtz resonator. The theo-
retical resonance criterions reasonably predicted
the occurrence of diffuser-flow pulsations.



NACA TN 3510

AN AUTOMATIC VISCOMETER FOR NON-
NEWTONIAN MATERIALS. Ruth N. ,Veltmann and
Perry W. Kuhns. August 1955. 34p. diagrs..
photos., tab. (NACA TN 3510)

A concentric-cylinder rotational viscometer is des-
cribed that can measure viscosities from 0.05 to
20,000 poises, program and record meaningful flow
curves of rate of shear against shearing stress for
most non-Newtonian materials, record time-torque
curves, and produce dynamic flow measurements.
A program feature provides for selecting (1) the
maximum rate of shear up to 4000 sec -1, (2) the
time required to increase rotational speed from zero
to the preset maximum, and (3) tNo sequences of
rate of shear, which is varied automatically at a
constant acceleration. Test data and curves are
presented.



NACA TN 3565

CHEMICAL ACTION OF HALOGENATED AGENTS
IN FIRE EXTINGUISHING. Frank E. Belles.
September 1955. 28p. diagrs., 3 tabs. (NACA
TN 3565)

The action of halogenated fire-fighting agents is dis-
cussed in terms of chain-breaking reactions between
agent and active particles (atoms and free radicals).
Literature data on flammability peaks of n-heptane-
agent-air mixtures are roughly correlated by reac-
tivities assigned to halogen and hydrogen atoms in
agent molecules. The value of hydrogen in agents is
discussed. A method for the estimation of rate con-
stants for the reaction of agent with active particles
is described, and the rate constant [or methyl bro-
mide is obtained. The result predicts the observed
peak concentration of methyl bromide quite well.
The assumed mechanism is consistent with the
observed facts.





NACA
RESEARCH ABSTRACTS NO. 89



BRITISH REPORTS


N-38122-

Royal Aircraft Establishment (Gt. Brit.)
THE REPRESENTATION OF ENGINE AIRFLOW IN
WIND TUNNEL MODEL TESTING. J. Seddon and
L. F. Nicholson. May 1955. 38p. diagrs., photos.
(RAE Tech. Note Aero 2371)

The problems of engine air-flow representation in
wind-tunnel models are reviewed. Methods which
have been used satisfactorily in low subsonic tunnels
are described briefly. Special dilficulties associ-
ated with testing at transonic speeds are noted.
Techniques ol special application to small supersonic
tunnels are described in some detail. It is shown
that there are reasons why the representation of jets
may be more important at supersonic speeds than at
subsonic speeds and a description is given of the
RAE jet interference tunnel, which is designed for
the study of some of the problems involved.


N-38123*

Royal Aircraft Establishment (Gt. Brit.)
BOUNDARY LAYERS ON SWEPT WINGS: THEIR
EFFECTS AND THEIR MEASUREMENT. D.
Kuchemann. April 1955. 37p. diagrs. (RAE
Tech. Note Aero 2370)

This is a brief review of some work which has been
done mainly at the RAE in the field of three-
dimensional boundary layers, in particular on swept
wings. Theoretical results are briefly mentioned,
and some examples of measured boundary layers are
given to illustrate the various effects that can now be
distinguished as being characteristic of swept wings.
The lift reduction and the drag increment which re-
sult from the boundary layer in comparison with the
values in inviscid flow are discussed; and some
types of flow which may occur as a consequence of
boundary-layer separation are described. In view
of the importance of obtaining further experimental
evidence, reference is made to the instrumentation
needed for that purpose.


N-38124'

Royal Aircraft Establishment (Gt. Brit.)
THE USE OF PITOT TUBES IN THE MEASURE-
MENT OF LAMINAR BOUNDARY LAYERS IN
SUPERSONIC FLOW. R. J. Monaghan. May 1955.
37p. diagrs., photos., 2 tabs. (RAE Tech.Note
Aero 2369)

This note summarizes the results of some research
on the laminar problem at the Royal Aircraft Estab-
lishment. It describes the various effects (real or
apparent) of pitot size found in a series of measure-
ments of laminar boundary layers in supersonic
flow. These lead to errors in the values of displace-
ment and momentum thickness calculated from the
pitot measurements, but it seems that it may be pos-
sible to obtain the correct experimental values by
applying a simple correction factor dependent only on
the ratio of pitot diameter to boundary-layer thick-
ness.


3


N-38466'

Royal Aircraft Establishment (Gt. Brit.)
PRESSURE-CABIN DESIGN: A DISCUSSION OF
SOME OF THE STRUCTURAL PROBLEMS IN-
VOLVED, WITH SUGGESTIONS FOR THEIR SOLU-
TION. D. Williams. March 1955. 51p. diagrs..
tab. (RAE Tech. Note Structures 155)

The structural factors governing the design of a
pressure cabin are discussed with a view to Lndi-
cating possible practical approaches.


N-38467"

Marine Aircraft Experimental Establishment
(Gt. Brit.) FULL SCALE MEASUREMENT OF
IMPACT LOADS ON A LARGE FLYING BOAT
(SUNDERLAND MK. 5). PART Im DATA FOR
IMPACTS ON MAIN STEP. R. Parker. August
1954. 37p. diagrs., 2 tabs. (MAEE F.'Res, 259)

The results of a series of full-scale impact tests on
the hull of a Sunderland Mk. 5 flying boat at an all
up weight of 50,000 lb were compared with an ap-
propriate theory and discussed generally in a pre-
vious report. As discrepancies between this data
and the theory were shown and explanation of this
has not been forthcoming, this report presents
complete time histories of a number of the actual
measurements to permit comparison with other
theories when such become available and application
by empirical methods where appropriate.


N-38468"

Marine Aircraft Experimental Establishment
(Gt. Brit.) INVESTIGATION OF HIGH LENGTH.
BEAM RATIO SEAPLANE HULLS WITH HIGH
BEAM LOADINGS HYDRODYNAMIC STABILITY.
PART 18 THE STABILITY AND SPRAY CHARAC-
TERISTICS OF MODEL N. D. M. Ridland, J. K.
Friswell, and A. G. Kurn. April 1955. 25p.
diagrs.. photos., 3 tabs. (MAEE F Res, 254)

In this report results are presented of limited tests
on the hydrodynamic characteristics of Model N of
the series, these tests being designed solely to pro-
vide information on the interactions of the different
parameters relevant to the series. The model has a
length-to-beam ratio of 13 (the forebody being 6
beams in length and the afterbody 7 beams), 40
forebody warp, an afterbody to forebody keel angle
of 80, and a straight transverse step with a step
depth of 0.15 beams. The tests comprised the
determination of longitudinal stability limits without
slipstream at Ca = 2.75 and an investigation of
spray at this loading. A short discussion of the
results is also included.


N-38469'

Marine Aircraft Experimental Establishment
(Gt. Brit.) INVESTIGATION OF HIGH LENGTH/
BEAM RATIO SEAPLANE HULLS WITH HIGH
BEAM LOADINGS HYDRODYNAMIC STABILITY.
PART 13 THE EFFECT OF AFTERBODY ANGLE
ON STABILITY AND SPRAY CHARACTERISTICS.
D. M. Ridland. February 1955. 31p. diagrs.,
photos., 3 tabs. (MAEE F/Res/248)









The effects of afterbody angle on longitudinal sta-
bility, spray and directional stability characteristics,
and elevator effectiveness are deduced from the re-
sults of tests on three models of the series which
were alike in every major respect except that of
afterbody angle. The models had afterbody angles of
40, 60, and 80, respectively. It was found that in-
creasing afterbody angle improved longitudinal sta-
bility characteristics considerably, both with and
without disturbance, increased trim generally, gave
a slight improvement in spray and directional sta-
bility characteristics, and increased elevator ef-
fectiveness. The best hydrodynamic configuration
was that with the 80 afterbody angle.

N-38470*

Royal Aircraft Establishment (Gt. Brit.)
THE USE OF NEGATIVE FEEDBACK FOR DE-
RIVING SHARP LIMITS IN ELECTRONIC ANA-
LOGUE COMPUTERS. M. Squires and P. R.
Benyon. April 1955. 27p. diagrs. (RAE Tech,
Note GW 361)

This note draws attention to improvements in elec-
tronic limiting action which results when standard
limiting devices are enclosed with d-c amplifiers
in negative feedback loops. Experiments in which a
"high-gain" amplifier was used yielded limiters
which were ideal to within practical limits of obser-
vation. Three types of limiter are recommended.
In one the limiting characteristic of an output cathode
follower forming part of the amplifier is used. In
the remaining two, the high reverse resistance of
diodes is used to give a single-sided and a two-sided
limit, respectively.


N-38471*

Royal Aircraft Establishment (Gt. Brit.)
HEAT TREATMENT AND TENSILE PROPERTIES
OF A TITANIUM 13-PERCENT MOLYBDENUM
ALLOY. G. I. Lewis, J. I. M. Forsyth, and
H. Brooks. March 1955. 10p. diagrs., photos.
(RAE Tech. Note Met. 215)

Titanium 13-percent molybdenum alloy strip
0.06 in. thick prepared from magnesium reduced
titanium sponge and molybdenum strip by arc melting
and hot and cold rolling had a strength of 48.6
tons/in. 2 with 33-percent uniform and 37-percent
total elongation after quenching from 8250 C to retain
the p phase. Strengths obtained by subsequent over-
aging in the a + P field at 5000-5500 C ranged from
79.3 tons/in.2 with 6-percent elongation and
E = 15.5 x 106 lb/in.2to 87.4 tons/in.2with4-percent
elongation and E = 15.9 x 106 lb/in.2. Deformation
in the quenched condition caused partial martensitic
breakdown of the metastable p structure and for the
one set of conditions investigated (test pieces
quenched from 8250 C, strained 10 percent in ten-
sion, and aged 24 hours at 5500 C) reduced the
strength and modulus obtained on aging. The den-
sity of the alloy was 4.92 gm/cc or 0.18 lb/cu in.


N-38472*

Royal Aircraft Establishment (Gt. Brit.)
SOME EXPERIMENTS ON THE EXTRUSION OF
MAGNESIUM AND ALUMINIUM POWDERS. H. G.
Cole. February 1955. 23p. diagrs., photos.,
5 tabs. (RAE Met. 85)


NACA
RESEARCH ABSTRACTS NO. 89

Smooth extrusions of good mechanical properties can
be made from magnesium powders at low extrusion
speeds provided the billet temperature exceeds
4700 C. Conditions for the extrusion of some alumi-
num powders coarser than S. A. P. have been deter-
mined. The relation between pressure and density
during the cold compacting of some magnesium and
aluminum powders is recorded. Remnants of natural
oxide films in extrusions often, but not always, stop
the propagation of slip during straining. Some
electron micrographs of aluminum oxide remnants
are shown.



N-38473*

Royal Aircraft Establishment (Gt. Brit.)
TRIALS OF BANNER TARGETS LAUNCHED ON A
FIXED TOW LENGTH. D. R. Bettison and
A. Eldridge. March 1955. 20p. diagrs., photos.,
3 tabs. (RAE Tech. Note Mech. Eng. 199)

As a result of flight trials it has been established
that a modified R.F.D. banner target can be launched
on a fixed tow length of 900 ft, at a speed of 130
knots and at altitudes up to 20, 000 ft. To avoid
tangling and searing of the lo* rope and safety
webbing link it is essential that, during the launch,
they are deployed from a tray fitted to the target
container, or in the aircraft bomb-bay.


N-38474*

Nat. Gas Turbine Establishment (Gt. Brit.)
THE PERFORMANCE OF A MULTI-STAGE COM-
BUSTION SYSTEM. PART I FLAME STABILITY
OF RECIRCULATION ZONES. S. W. White and
W. G. E. Lewis. May 1955. 16p. diagrs.
(NGTE R.165)

As part of a general investigation into optimum rates
of addition of combustible mixtures into a gaseous
zone at high temperature, this report describes an
experimental study into the factors influencing flame
stability of cylindrical baffles admitting mixtures
through a single row of holes in the periphery, the
axis of the cylinder being parallel to the approaching
airstream. The results show that for geometrically
similar baffles the stability range can be correlated
with Q where Q is the air flow rate, D
D2. 4p1.5
is the baffle diameter and P is the combustion
pressure.



N-38499*

Royal Aircraft Establishment (Gt. Brit.)
THE FLOW THROUGH SHORT STRAIGHT PIPES IN
A COMPRESSIBLE VISCOUS STREAM. J. Seddon.
April 1955. 34p. diagrs., photos. (RAE Aero 2542)
In the design of small models of aircraft or missiles
for testing in supersonic tunnels, it may be desired
to represent jet engine nacelles by means of simple
hollow pipes. The note sets out the principal char-
acteristics of compressible flow in such pipes at
zero yaw and gives a theory for calculating the effect
of the boundary layer. This is checked against the
results of tests with a series of pipes of varying






NACA
RESEARCH ABSTRACTS NO. 89


size, at Mach numbers from 1.34 to 2.41. Curves
are presented for determining the maximum length
radius ratio of a parallel pipe which will permit
supersonic internal flow, in terms of the Mach num-
ber of the stream and Reynolds number of the pipe:
the curves are given for both laminar and turbulent
internal boundary layers. The effect of inclination
of the pipe to the stream is discussed briefly, on the
basis of results at one Mach number (1.86).



N-38501'

Royal Aircraft Establishment (Gt. Brit.)
PRELIMINARY LOW SPEED WIND TUNNEL TESTS
ON FLAT PLATES AND AIR BRAKES: FLOW,
VIBRATION AND BALANCE MEASUREMENTS.
R. Fail, T. B. Owen. and R. C. W. Eyre. January
1955. 45p. diagrs., 5 tabs. (RAE Tech. Note
Aero 2356)

Flow measurements have been made behind sharp-
edged flat plates: (a) at 900 for various shapes, and
(b) for a square plate over an incidence range. The
results of (a) show a closed bubble about 3 plate
sides long, with a constant pressure boundary up to
the maximum diameter, followed by mixing. Meas-
urements of velocity fluctuations were made for (b),
showing that a regular shedding of turbulent eddies
occurs for 0 = 500 and over, but stops by 400.
Large random low frequency longitudinal fluctuations
are associated with the shedding. Lift, drag, and
pitching moment increments were measured on a
square plate motuited on (a) a long cylinder, and (b)
near the end of several shapes ol rear fuselage, to
see how the moments on opening the brake could be
modified. Velocity fluctuation measurements, at
700 only, show a much reduced longitudinal un-
steadiness when the plate is in proximity to the
fuselage. Comparative rests were made on a cas-
cade brake and show that no shedding occurs.


N-38510'

Royal Aircraft Establishment (Gt. Brit.)
THE INFLUENCE OF METHYL BROMIDE ON
FLAMES. I PREMIXED FLAMES. 1 -
%DIFFUSION-FLAMES. R. F. Simmons and H. G.
Wolfhard. March 1955. 27p. diagrs., photos.,
3 tabs. (RAE RPD 25)

The influence of methyl bromide and bromine on
flames has been investigated. Both premixed and
diffusion flames are chemically inhibited, in con-
trast to the purely diluting effect of inert gas. The
most significant change in premixed flames on add-
ing methyl bromide or bromine is that the limit
temperature is raised. Limits of inflammability of
mixtures cannot be calculated by Le ChatelLer's
rule. The action of methyl bromide in flames can be
interpreted in terms of the addition of equivalent
amounts of fuel and bromine. The structure of diflu-
sion flames is completely changed when methyl bro-
mide is introduced into the air, an additional reac-
tion zone being formed on the air side in which the
methyl bromide reacts with the air. Thus fuel and
oxidant no longer meet in simple stoichiometric pro-
portions in the main reaction zone. Methyl bromide
has been found to be much more efficient as anextin-
guishant when introduced into the air than when
added.to the fuel.


5



N-38511'

Royal Aircraft Establishment (Gt. Brit.)
AN ELECTRONIC SINE WAVE GENERATOR FOR
FREQUENCIES DOWN TO 1/100 C. P.S. O. H.
Lange, J. P. Lonergan, and J. J. Gait. March
1955. 14p. diagrs. (RAE TECH. Note GW 360)

This paper describes an electronic generator of low
output impedance for the production of sine waves of
controlled amplitude in the frequency band from
1 '100 cps to 10 cps.


N-38517'

Royal Aircraft Establishment (Gt. Brit.)
A FIXTURE FOR TESTING SHEET MATERIALS IN
COMPRESSION AT ELEVATED TEMPERATURES.
D. C. Hayward. March 1955. 19p. diagrs., photos.,
3 tabs. (RAE Tech. Note Met. 193)

A fixture for compressive testing of sheet material at
elevated temperatures up to 4000 C has been devel-
oped. The thinnest sheet tested was 20 S. W. G.
(0.036 in.) Lateral support, to prevent buckling of
the test specimens under high stress, was obtained
from a series of steel balls spaced evenly on both
sides of the specimen. An averaging extensometer,
attached to opposite edges of the specimen, recorded
changes in gage length and enabled readings of strain
to include the 0.5 percent proof stress to be made.
Compressive test readings compared favorably both
for sensitivity and reproducibility with corresponding
readings taken by established tensile test methods,
the facility and simplicity of the compression test
procedure equaling that of the tensile lest.


N-38520'

Royal Aircraft Establishment (Gt. Brit.)
STATIONARY FLAMES WITH METHYL NITRATE
AND METHYL NITRITE. P. Gray, A. R. Hall,
and H. G. Wolfhard. March 1955. 24p. diagrs.,
photos. (RAE RPD 26)

A kinetic and spectroscopic study was made at low
pressures of a stationary, self-decomposition
methyl nitrate flame, premixed with oxygen, nitric
oxide, and nitrogen dioxide. Another flame,
methyl nitrite and oxygen, was studied and was
found to be similar in appearance and spectra at
similar conditions. The basic features of these
flames have been interpreted in terms of the oc-
currence of pyrolysis reactions before the main
flame zone is reached and the stability of nitric
oxide at lower temperatures and pressures.


N-38521*

Royal Aircraft Establishment (Gt. Brit.)
CORROSION FATIGUE TESTS AT SLOW RATES OF
LOADING ON ALUMINIUM ALLOY SHEET TO
SPEC. D.T.D. 546. J. T. Ballett and M. S.
Binning. 9p. diagrs., photo., 2 tabs. (RAE Tech.
Memo. Met. 145 (Second Issue))

Comparative fatigue tests were made on simple
specimens of a De Havilland Comet fuselage skin
material in air and in water. A testing machine
was made which applied a loading the same as that









used to pressurize the fuselage in a full-scale test.
The results showed that in repeated load fatigue
tests for endurances of up to about 5000 cycles at a
frequency of 1 cycle in 5 minutes, the endurances
are not appreciably affected by immersing the speci-
mens in water during testing.


N-38532*

Royal Aircraft Establishment (Gt. Brit.)
SOME CREEP PROPERTIES OF NIMONIC 95 AT
8500 C AND NIMONIC 80A (SPECIAL PROCESS) AT
7500 C AND 8150 C. L. W. Larke and R. A.
Whittaker. March 1955. 23p. diagrs., 7 tabs.
(RAE Tech. Note Met. 217)

The overall purpose of the program is to determine
the stresses necessary to produce total plastic
strains of 0.1 percent to 1.0 percent and fracture
where possible in times up to 2, 000 hours at tem-
peratures of 8500, 8750, and 9000 C on Nimonic 95
and at 7500, 8150, and 8750 C on Nimonic 80A
"special process" material. On Nimonic 95 at
8500 C, 5t/in. 2 produces 0. 1 percent strain in
1, 000 hours. The data on Nimonic 80A "special
process" for 0.1 percent strain in 1,000 hours are
4. 9t/in. 2 at 7500 C and 1. 7t/in. 2 at 8150 C.
Comparison is made of test results for the two
materials.


N-38533*

Royal Aircraft Establishment (Gt. Brit.)
FURTHER WORK ON THE EFFECT OF HARD-
ANODISING ON THE FATIGUE STRENGTH OF
D. T.D.364B ALUMINIUM ALLOY. E. G. Savage
and E. G. F. Sampson. March 1955.
lip. diagrs., 5 tabs. (RAE Tech. Note Met. 216)

As suggested in a previous note, a modified form of
treatment was used to produce hard anodic coatings
on Wohler-type fatigue test-pieces made from
D. T. D. 364B aluminum alloy, with the object of
reducing the ill effects due to treatment at a constant
current density. This modified treatment failed to
reduce the fatigue loss significantly. However, it
was found that sealing the coatings in a boiling
dichromate solution substantially improved the fatigue
strength of parts which had been hard anodised by
either the "standard" or the modified process, the
losses of fatigue strength at 10 million cycles being
reduced from 47 to 8 percent and 36 to 4 percent,
respectively. The sealing treatment reduced the
abrasion-resistance of the coatings by 20 percent.
No further work is proposed.



N-38559*

Royal Aircraft Establishment (Gt. Brit.)
DIODE TEST EQUIPMENT TYPE NO. 2281 AS
MODIFIED FOR THE TESTING OF VALVES, TYPE
VX.9156. J. C. Le Grice. April 1955. 16p.
diagrs., photos. (RAE Tech. Note EL.90)

Operating instruction, circuit descriptions and
calibration procedure are given for this equipment
which performs striking voltage measurements on
VX.9156 cold cathode diodes as required by their
acceptance specification.


NACA
RESEARCH ABSTRACTS NO. 89


N-38561*

Royal Aircraft Establishment (Gt. Brit.)
A COMPARISON OF SOME DEVICES FOR REDUCING
THE GROUND RUN DURING LANDING. D. Lean.
May 1955. 29p. diagrs. (RAE Tech. Note Aero 2379)

With the thrust/weight ratios common today on jet
aircraft, the size of brake parachute needed to equal
the effect of the available reversed thrust is large
but not excessive. For the future, increased touch-
down speeds tend to favor the brake parachute,
simply by increasing the mean drag. This gain is
probably only partly cancelled by higher available
reversed thrust. On the score of weight, cost, and
complications, the parachute has the advantage, but
from the handling point of view, pilots are likely to
prefer the more controllable reversed thrust system.
Too little is yet known of actual installations of
thrust reversal schemes on typical aircraft for a
comparison to be made on the grounds of reliability,
or effect on the aircraft and its occupants.

N-38562*

Royal Aircraft Establishment (Gt. Brit.)
FLIGHT EXPERIMENTS ON BOUNDARY LAYER
CONTROL FOR LOW DRAG. M. R. Head,
D. Johnson, and M. Coxon. March 1955. 48p.
diagrs., photos., 2 tabs. (RAE Aero 2541)

Tests have been made with distributed suction
applied to a short span sleeve fitted to the upper
surface of the wing of a single seat Vampire aircraft.
Full chord laminar flow was maintained up to
Reynolds numbers in the region of 29 million and
Mach numbers up to 0.70, which was very nearly the
critical Mach number of the sleeve section. The
suction quantities required were sufficiently small
to result in overall reductions in profile drag of be-
tween 70 and 80 percent, account being taken of the
power required for suction. Difficulties were ex-
perienced due to surface roughness, and although
these are believed to have resulted largely from the
particular type of porous covering used in the tests,
the problem of maintaining a sufficiently smooth and
clean surface is evidently of crucial importance to
full-scale application.


N-38563*

Royal Aircraft Establishment (Gt. Brit.)
STRESS CONSIDERATIONS IN THE DESIGN OF
PRESSURIZED SHELLS. E. H. Mansfield. April
1955. 27p. (RAE Structures 178)

This report considers the design and stress analysis
of pressurized thin-walled shells with special refer-
ence to openings in the shell wall.


N-38565*

Royal Aircraft Establishment (Gt. Brit.)
OPTICAL CHARACTERISTICS OF LAMINATED
CAMERA WINDOWS. A. C. Marchant and B. M.
Mathieson. March 1955. 24p. diagrs., photos.,
2 tabs. (RAE Tech.Note Ph.489)

This note describes optical tests on laminated
camera windows of both plane and spherical form.
Photographic tests of their effect upon the resolving






NACA
RESEARCH ABSTRACTS NO. 89

power of an associated camera system have been
carried out. The wedge angle of the windows has
been measured, and the surface flatness and internal
homogeneity of the plane windows tested; while the
effects of the curved windows, upon the correction of
the camera lens associated with them, has been as-
sessed visually. It is concluded that the laminating
process does not, in itself, introduce deleterious ef-
fects on optical performance. Plane windows, unless
optically worked to a high degree of flatness, may
introduce a focal shift which will, in practice, reduce
resolving power. Curved windows, even when
worked to have zero power, are shown to reduce
overall resolution through the introduction of astig-
matism and field curvature.


N-38609*

Royal Aircraft Establishment (Gt. Brit.)
IMPURITIES IN TITANIUM: SULPHUR. D. A.
Sutcliffe. March 1955. 13p. diagrs., photos., 7
tabs. (RAE Tech. Note Met. 218)

Binary titanium sulphur alloys contairung up to 1.05
percent sulphur have been prepared by arc melting.
Tensile, hardness, and impact tests have been made
on the alloys at room temperature. The results
show that the addition of approximately 0.1 percent
sulphur causes a marked rise in tensile strength and
hardness with a corresponding fall in elongation and
impact strength. Further additions up to 1.05 per-
cent sulphur cause only a small increase in tensile
strength and hardness, but the elongation and impact
strength continue to decrease steadily, though less
rapidly than at first.


MISCELLANEOUS


N-37766"

THE ICING PROBLEM: CURRENT STATUS OF
NACA TECHNIQUES AND RESEARCH. Uwe H.
von Glahn. (Preprint of paper presented at Ottawa
AGARD Conference, June 10-17, 1955). 44p.
diagrs., photos.

Icing tunnel facilities, operational techniques, and
icing instruments are described in appendices to
this paper.


DECLASSIFIED NACA REPORTS


NACA RM A50L04

PRELIMINARY FLIGHT INVESTIGATION OF THE
MANEUVERING ACCELERATIONS AND BUFFET
BOUNDARY OF A 350 SWEPT-bVING AIRPLANE AT
HIGH ALTITUDE AND TRANSONIC SPEEDS.
George A. Rathert, Jr., Howard L. Ziff, and George
E. Cooper. February 21, 1951. 12p. diagrs.,
photo., tab. (NACA RM A50LO4) (Declassified
from Confidential, 7/20, 55)

Results are presented from a series of exploratory
flights on a swept-wing fighter airplane up to 1.09
Mach number to show the effects of compressibility
presently limiting the maneuverability. The buffet
boundary is also presented.


7

NACA RM A51B21

TESTS IN THE AMES 40- BY 80-FOOT WIND TUN-
NEL OF AN AIRPLANE CONFIGURATION WITH AN
ASPECT RATIO 2 TRIANGULAR WING AND AN
ALL-MOVABLE HORIZONTAL TAIL-
LONGITUDINAL CHARACTERISTICS. David Graham
and David G. Koenig. April 23, 1951. 35p. diagrs.,
photo., 3 labs. (NACA RM A51B21) (Declassified
from Confidential, 8/17,55)

Force tests were made with an all-movable horizon-
tal tall in three vertical positions above the extended
wing-chord plane (0, 0.25, and 0.50 wing semispan)
at one longitudinal distance aft of the wing. Down-
wash variations at the position of the horizontal tail
as obtained from force tests and from a downwash
survey are compared. The results indicated that,
from a standpoint of longitudinal stability at low
speed, the best of the positions of the horizontal tail
tested would be that in the extended wing-chord
plane. The Reynolds number of the tests was
14.6 x 106 and the Mach number was 0.13.


NACA RM A51HIOa

TESTS IN THE AMES 40- BY 80-FOOT WIND TUN-
NEL OF AN AIRPLANE CONFIGURATION WITH AN
ASPECT RATIO 4 TRIANGULAR WING AND AN
ALL-MOVABLE HORIZONTAL TAIL LONGITUDI-
NAL CHARACTERISTICS. David Graham and
David G. Koenig. October 1951. 27p. diagrs.,
photo., 4 tabs. (NACA RM A51H0a) (Declassified
from Confidential, 8'17,'55)

Force tests were made of the wing alone, wing fuse-
lage, wing-fuselage vertical-tail, and complete con-
figuration with each of two horizontal tails. The
tails were located at each of three vertical positions
above the extended wing-chord plane (0, 0.18, and
0.36 wing semispan) at one longitudinal distance aft
of the wing. Downwash variations as obtained from
force tests are presented. The Reynolds number of
the tests was 10.9 x 106 and the Mach number was
0.13. The results indicated that, from a standpoint
of longitudinal stability at low speed, the best of the
positions of the horizontal tail tested would be that
in the extended wing-chord plane.

NACA RM A53E01

COMPARISON OF THE EXPERIMENTAL AND
THEORETICAL DISTRIBUTIONS OF LIFT ON A
SLENDER INCLINED BODY OF REVOLUTION AT
M = 2. Edward W. Perkins and Donald M. Kuehn.
August 1953. 39p. diagrs., photos., 2 tabs. (NACA
RM A53E01) (Declassified from Confidential,
8/17/55)

An investigation of the pressure distribution for a
body of revolution, consisting of a 33-1/3-calibei
(5.75 fineness ratio) tangent ogival nose and a
cylindrical afterbody, has been made for an angle-
of-attack range of 00 to 35.50 at a Mach number of
1.98 and a Reynolds number of approximately
0.5 x 106, based on body diameter. Comparisons of
the theoretical and experimental pressure distribu-
tions are made to show the nature of the effects of
both viscosity and cross-flow compressibility. The
experimental load distributions are compared with
those predicted by the method suggested in NACA
RM A9126 which includes an approximate method for
taking into account the effects of viscosity on the
lift distribution.





8



NACA RM E50H07

PERFORMANCE OF 4600-POUND-THRUST
CENTRIFUGAL-FLOW-TYPE TURBOJET ENGINE
WITH WATER-ALCOHOL INJECTION AT INLET.
Philip W. Glasser. October 9, 1950. 24p. diagrs.
(NACA RM E50H07) (Declassified from Confidential,
8/17/55)

An experimental investigation of the effects of in-
jecting a water-alcohol mixture of 2:1 at the com-
pressor inlet of a centrifugal-flow-type turbojet
engine was conducted in an altitude test chamber at
static sea-level conditions and at an altitude of
20,000 feet with a flight Mach number of 0.78 with an
engine operating at rated speed. The net thrust was
augmented by 0.16 for both flight conditions with a
ratio of injected liquid to air flow of 0.05. Further
increases in the liquid-air ratio did not give compa-
rable increases in thrust.




NACA RM E50H30

ALTITUDE INVESTIGATION OF PERFORMANCE OF
TURBINE-PROPELLER ENGINE AND ITS COMPO-
NENTS. Lewis E. Wallner and Martin J. Saari.
October 5, 1950. 65p. diagrs., photos. (NACA
RM E50H30) (Declassified from Confidential,
8/17/55)

An investigation was conducted on a turbine pro-
peller engine in the NACA Lewis altitude wind tunnel
at altitudes from 5,000 to 35,000 feet. The applica-
bility of generalized parameters to turbine-propeller
engine data, analyses of the compressor, the com-
bustion chambers, and the turbine, and a study of
the overall engine performance are reported.
Engine performance data obtained at sea-level static
conditions could be used to predict static perform-
ance at altitudes up to 35,000 feet by use of the
standard generalized parameters.




NACA RM E52H07

PROCEDURE FOR CALCULATING TURBINE
BLADE TEMPERATURES AND COMPARISON OF
CALCULATED WITH OBSERVED VALUES FOR
TWO STATIONARY AIR-COOLED BLADES. W.
Byron Brown, Henry O. Slone, and Hadley T.
Richards. September 1952. 38p. diagrs. (NACA
RM E52H07) (Declassified from Confidential,
8/17/55)

Local and average blade temperatures were calcu-
lated for two stationary air-cooled turbine blades
with 10 tubes and 13 fins forming the internal heat-
transfer surfaces. These temperatures were calcu-
lated using previously published NACA temperature-
distribution equations and the most recent theories
for determining heat-transfer coefficients, including
for the first time the allowance for effects of vari-
able wall temperature on gas-to-blade heat-transfer
coefficients at the leading and trailing sections of
turbine blades. Comparison of calculated and ex-
perimental blade temperatures, for gas tempera-
tures of 3000 and 10000 F, resulted in good agree-
ment.


NACA
RESEARCH ABSTRACTS NO. 89


NACA RM E53H05

CONSIDERATIONS IN THE ADAPTATION OF LOW-
COST FUELS TO GAS-TURBINE-POWERED COM-
MERCIAL AIRCRAFT. Henry C. Barnett and
Richard J. McCafferty. October 1953. 59p.
diagrs., photo., 2 tabs. (NACA RM E53H05)
(Declassified from Confidential, 8/17/55)

In recent months interest has increased in the possi-
ble use of distillate and residual fuel oils as fuels for
commercial gas-turbine aircraft. However, the use
of such fuels entails the solution of many problems
pertaining to fuel physical properties and combustion
characteristics. This report reviews some of these
problems and discusses the status of current knowl-
edge in relation to their solution. The fuel combus-
tion characteristics considered in this study are
ignition, combustion efficiency, combustion stability,
and carbon deposition. In addition the effects of
distillate and residual fuels on exhaust deposits and
corrosion are discussed. Preliminary calculated
data are presented to indicate preheating require-
ments for maintaining fuel fluidity and desirable
fuel-injection properties.




NACA RM L7H26

FLIGHT INVESTIGATION TO DETERMINE THE
HINGE MOMENTS OF A BEVELED-EDGE AILERON
ON A 450 SWEPTBACK WING AT TRANSONIC AND
LOW SUPERSONIC SPEEDS. William N. Gardner
and Howard J. Curfman, Jr. November 12, 1947.
20p. diagrs., photos. (NACA RM L7H26)
(Declassified from Confidential, 8/17/55)

An RM-1 stability and control research pilotless air-
craft was used to measure aileron hinge moments in
flight. The 200 beveled edge, 0.52ca overhang
balance aileron experienced a rapid increase and then
decrease in hinge moments as the critical speed was
exceeded (M = 0.86 to 0.92). At supercritical speeds
(M = 0.92 to 1.23) the hinge moments increased
rapidly. Rolling-moment balance data indicate
aileron effectiveness reversal at supercritical
speeds.



NACA RM L8E04

EFFECT OF WINDSHIELD SHAPE OF A PILOT'S
CANOPY ON THE DRAG OF AN NACA RM-2 DRAG
RESEARCH MODEL IN FLIGHT AT TRANSONIC
SPEEDS. Sidney R. Alexander. July 21, 1948.
6p. diagrs. (NACA RM L8E04) (Declassified from
Confidential, 8/17/55)

Results of flight tests of an NACA RM-2 drag re-
search model equipped with a pilot's canopy having
a vee windshield are presented for a Mach number
range from 0.75 to 1.43. Comparison is made with
test results of a similar canopy having a flat wind-
shield. The vee-windshield canopy produced lower
drag-coefficient values than the flat -windslueld
canopy for Mach numbers from 0.85 to about 1.2.
From M = 1.2 to 1.4 both canopies produced the
same drag coefficient.





NACA
RESEARCH ABSTRACTS NO. 89


NACA RM L8E10

FREE-FLIGHT INVESTIGATION OF THE ROLLING
EFFECTIVENESS AT HIGH SUBSONIC, TRANSONIC,
AND SUPERSONIC SPEEDS OF LEADING-EDGE
AND TRAILING-EDGE AILERONS IN CONJUNCTION
WITH TAPERED AND UNTAPERED PLAN FORMS.
H. Kurt Strass. July 23, 1948. 19p. diagrs.,
photos. (NACA RM L8E10) iDeclassflied from
Confidential, 8 17 '55)

Additional results of an aerodynamic-control-
effectiveness investigation using free-flight, rocket-
propelled test vehicles have been obtained recently
which show some effects of leading-edge and trailing-
edge ailerons tested both individually and in combi-
nation on wings with tapered and untapered plan
forms.


NACA RM L8E12

LONGITUDINAL STABILITY CHARACTERISTICS OF
A 420 SWEPTBACK WING AND TAIL COMBINATION
AT A REYNOLDS NUMBER OF 6.8 x 106. Stanley H.
Spooner and Albert P. Martina. July 22, 1948.
44p. diagrs., photos., 2 tabs. (NACA RM L8E12)
(Declassified from Confidential, 8 17. 55)

Presents results of wind-tunnel investigation at a
Reynolds number of 6.8 x 106 to determine the static
longitudinal stability characteristics of a 420 swept-
back wing and fuselage combination with a swept-
back horizontal tail. Includes effects of vertical
position of fuselage and tall with respect to wing for
several combinations of high-lift and stall-control
devices. Also includes effect of a simulated ground.

NACA RM L8G20a

FREE-FLIGHT INVESTIGATION AT TRANSONIC
AND SUPERSONIC SPEEDS OF THE ROLLING EF-
FECTIVENESS OF A THIN, LTNSWvEPT WING
HAVING PARTIAL-SPAN AILERONS. Carl A.
Sandahl. Cctober 22, 1948. 13p. diagrs., photos.,
tab. (NACA RM L8G20al (Declassified from
Confidential, 7 20, 55)

Tests of the rolling effectiveness at transonic and
supersonic speeds of a thin, unswept wing having
partial-span ailerons have been made by means of
rocket-propelled test vehicles. The results showed
that with 4.60 aileron deflection, the only deflection
tested, the rolling effectiveness decreased abruptly
in the Mach number range from about 0.92 to 0.97.
At supersonic speeds the rolling effectiveness was
considerably lower than at subsonic speeds. The
direction of roll was, in the correct sense, over the
entire Mach number range investigated.


NACA RM L9H19

A STUDY OF SEVERAL FACTORS AFFECTING THE
STABILITY CONTRIBUTED BY A HORIZONTAL
TAIL AT VARIOUS VERTICAL POSITIONS ON A
SWEPTBACK-WING AIRPLANE MODEL. Gerald V.
Foster and Roland F. Griner. October 28, 1949.
28p. diagrs., tab. (NACA RM L9H19) (Declassi-
fied from Confidential, 8/17/55)

Results are presented of a study made in the Langley
19-loot pressure tunnel at a Reynolds number of


9


6.8 x 106 to determine the effects of fuselage after-
body shape, split flaps, and a reduction in the span
of 0.575-span leading-edge flaps on the stability con-
tributed by a horizontal tall at various heights to a
420 sweptback wing-fuselage combination. Results
are also presented of air-flow surveys in the vicinity
of the tail with the wing flaps deflected.


NACA RM LSOGl4b

EXPERIMENTAL DETERMINATION OF EFFECT
OF STRUCTURAL RIGIDITY ON ROLLING EFFEC-
TIVENESS OF SOME STRAIGHT AND SWEPT WINGS
AT MACH NUMBERS FROM 0.7 TO 1.7. H. Kurt
Strass, E. M. Fields, and Paul E. Purser.
October 4, 1950. 40p. diagrs., photo., tab. (NACA
RM L50G14b) (Declassified from Confidential,
7. 20 55)

The effect of varying the wing structural rigidity on
the control effectiveness of a 0.2-chord, plain,
faired, full-span aileron on untapered wing plan
forms having 00 and 450 sweep and aspect ratio of
3.7 has been investigated with rocket-propelled test
vehicles. The test results are compared with results
obtained from existing theory. Aerodynamic twisting
moments are evaluated from the experimental
rolling-loss data.

NACA RM L50H28

TABULATED PRESSURE COEFFICIENTS AND
AERODYNAMIC CHARACTERISTICS MEASURED ON
THE WING OF THE BELL X-l AIRPLANE IN PULL-
UPS AT MACH NUMBERS FROM 0.53 TO 0.99.
Ronald J. Knapp and Gertrude V. Wilken.
November 1, 1950. 77p. diagrs., photo., 11 tabs.
(NACA RM L50H28) (Declassified from Confidential,
8/17/55)

Presents tabulated pressure coefficients and aero-
dynamic characteristics measured on the left wing
of the Bell X-l airplane in 10 pull-ups from Mach
numbers of approximately 0.53 to 0.99.

NACA RM L50J16

DYNAMIC STABILITY AND CONTROL CHARACTER-
ISTICS OF A VERTICALLY RISING AIRPLANE
MODEL IN HOVERING FLIGHT. William R. Bates,
Powell M. Lovell, Jr., and Charles C. Smith, Jr.
February 23, 1951. 16p. diagrs., photos., tab.
(NACA RM L50JI6) (Declassified from Confidential,
8/17/55)

An investigation is being made to determine the sta-
bility and control characteristics of a vertically
rising airplane model. This paper presents the re-
sults of some preliminary hovering flight tests
made in still air with two center-of-gravity positions,
0-percent and 45-percent mean aerodynamic chord
and with normal airplane-type controls operating in
the slipstream.

NACA RM L51G05

EFFECT OF THE PROXIMITY OF THE GROUND ON
THE STABILITY AND CONTROL CHARACTERIS-
TICS OF A VERTICALLY RISING AIRPLANE MODEL
IN THE HOVERING CONDITION. Charles C. Smith,
Jr., Powell M. Lovell, Jr., and William R. Bates.
September 1951. 16p. diagrs., tab. (NACA
RM L51G05) (Declassified from Confidential,
8/17/55)









An investigation has been made to determine the ef-
fect of the proximity of the ground on the stability
and control characteristics of a vertically rising air-
plane model in the hovering condition. The model
was essentially a conventional airplane model having
a dual-rotating propeller in a tractor arrangement, a
rectangular wing and a cruciform tail with rectangu-
lar surfaces. The investigation included take-offs
and landings and hovering flight tests near the
ground, force tests with and without a ground board,
and dynamic-pressure surveys at various radial and
longitudinal stations behind the propeller.


NACA RM L51H13a

DYNAMIC STABILITY AND CONTROL CHARACTER-
ISTICS OF A DELTA-WING VERTICALLY RISING
AIRPLANE MODEL IN TAKE-OFFS, LANDINGS,
AND HOVERING FLIGHT. Powell M. Lovell, Jr.,
William R. Bates and Charles C. Smith, Jr.
October 1951. 14p. diagrs., photo., tab. (NACA
RM L51H13a) (Declassified from Confidential,
8/17/55)

An investigation has been conducted by means of
tests of a flying model in still air to determine the
dynamic stability and control characteristics of a
delta-wing vertically rising airplane in the take-off,
landing, and hovering phases of flight. The model
had a dual-rotating propeller in a tractor arrange-
ment, a modified triangular wing, and modified tri-
angular vertical-tail surfaces mounted symmetrically
above and below the fuselage, but had no horizontal
tail. Control was provided by elevons and rudders
operating in the propeller slipstream.

NACA RM L51H14

DAMPING IN ROLL OF STRAIGHT AND 450 SWEPT
WINGS OF VARIOUS TAPER RATIOS DETERMINED
AT HIGH SUBSONIC, TRANSONIC, AND SUPER-
SONIC SPEEDS WITH ROCKET-POWERED
MODELS. E. Claude Sanders, Jr. October 1951.
15p. diagrs. (NACA RM L51H14) (Declassified
from Confidential, 8/17/55)

Rocket-powered flight tests have been conducted to
determine the damping in roll of several wings of 00
and 450 quarter-chord line sweep with various taper
ratios. The Mach number range of these tests was
from 0.8 to 1.45. Damping in roll decreased with
decreasing taper ratio at approximately the same
rate for swept and unswept wings, and was also de-
creased by sweeping the quarter-chord line. Exper-
imental data were much lower than that predicted by
theory for the swept wings. The drag at zero lift was
consistently of a lower magnitude for the 450 swept
tapered series than for the unswept tapered series.

NACA RM L51H16

INVESTIGATION OF THE EFFECT OF A NACELLE
AT VARIOUS CHORDWISE AND VERTICAL POSI-
TIONS ON THE AERODYNAMIC CHARACTERISTICS
AT HIGH SUBSONIC SPEEDS OF A 450 SWEPTBACK
WING WITH AND WITHOUT A FUSELAGE.
H. Norman Silvers, Thomas J. King, Jr., and
Thomas B. Pasteur, Jr. September 1951. 71p.
diagrs., photos., 3 tabs. (NACA RM L51H16)
(Declassified from Confidential. 8/17/55)


NACA
RESEARCH ABSTRACTS NO. 89


An investigation of a nacelle was made over a Mach
number range from 0.40 to 0.90 to determine the in-
terference characteristics between the nacelle and a
semispan model having a wing with the quarter-
chord line swept back 450 and a fuselage of fineness
ratio 10. The nacelle was a body of revolution of
fineness ratio 5.0 with a profile that was a modified
NACA 65-series airfoil section. The nacelle was
investigated in several chordwise positions and
vertical locations.

NACA RM L51I07a

ADDITIONAL STUDIES OF THE STABILITY AND
CONTROLLABILITY OF AN UNSWEPT-WING
VERTICALLY RISING AIRPLANE MODEL IN
HOVERING FLIGHT INCLUDING STUDIES OF VAR-
IOUS TETHERED LANDING TECHNIQUES.
William R. Bates, Powell M. Lovell, Jr., and
Charles C. Smith, Jr. November 1951. 25p.
diagrs., photos., tab. (NACA RM L51107a)
(Declassified from Confidential, 8 17/55)

This paper is the third of a series presenting the re-
sults of an investigation of the stability and control
characteristics of a flying model of an unswept-wing
vertically rising airplane. The investigation con-
sisted of flight tests to determine the effects of some
miscellaneous factors on the stability and control
characteristics for the hovering condition and to
study the behavior of the model for various landing
techniques involving the use of lines for pulling
the model in for a landing.

NACA RM L52H12

LOW-LIFT BUFFET CHARACTERISTICS OB-
TAINED FROM FLIGHT TESTS OF UNSWEPT THIN
INTERSECTING SURFACES AND OF THICK 350
SWEPTBACK SURFACES. Homer P. Mason.
January 1953. 21p. diagrs., photos. (NACA
RM L52H12) (Declassified from Confidential,
8/17/55)

Results of flight tests of two rocket-propelled re-
rearch models are presented. The effect of the in-
tersection of thin surfaces and the effect of moderate
sweepback of a thick surface on the lo-lift buffeting
and trim characteristics of a clean body-tail con-
figuration are shown. Drag coefficients are pre-
sented and compared with previous data to show the
effect of tail position and sweep.

NACA RM L53G15

INVESTIGATION OF A RELATED SERIES OF
TURBINE-BLADE PROFILES IN CASCADE. James
C. Dunavant and John R. Erwin. December 1953.
100p. diagrs. (NACA RM L53G15) (Declassdied
from Confidential, 8/17/55)

Series of directly related turbine blade shapes unm-
formly variable in camber are designed to satisfy
the usual turbine-blade operating requirements of
inlet air angle and turning angle. Low-speed cas-
cade test results of five blade sections and sample
high-speed cascade tests for two blade sections of
these series are presented. A method of properly
cambering turbine blade shapes based on low -speed
test results is given.





NACA
RESEARCH ABSTRACTS NO. 89


NACA RM L53H21

INVESTIGATION OF REYNOLDS NUMBER EF-
FECTS FOR A SERIES OF CONE-CYLINDER
BODIES AT MACH NUMBERS OF 1.62, 1.93. AND
2.41. Carl E. Grigsby and Edmund L. Ogburn.
October 1953. 20p. diagrs., photos. (NACA
RM L53H21) (Declassified from Confidential,
8/17. 55)

An investigation of the Reynolds number for transi-
tion and the skin-irlction drag at zero lift of eight
cone-cylinder bodies having varying fineness ratios
has been made at Mach numbers of 1.62, 1.93, and
2.41 over a Reynolds number range from 0.3 x 106 to
10 x 106. The accuracy of the skin-friction data was
not sufficient to permit any general conclusions to be
drawn. The Reynolds number for transition was
found to be dependent upon both the tunnel stagnation
pressure and Mach number.








NACA RM L53K12

A SUMMARY OF INFORMATION ON SUPPORT
INTERFERENCE AT TRANSONIC AND SUPERSONIC
SPEEDS. Eugene S. Love. January 1954. 26p.
diagrs. INACA RM L53K12) (Declassified from
Confidential, 8. 17. 55)

A compilation of available information on the prob-
lem of support interference at transonic and super-
sonic speeds is presented.


NACA RM L53L23a

WIND-TUNNEL INVESTIGATION OF A SHIELDED
TOTAL-PRESSURE TUBE AT A MACH NUMBER OF
1.61. Walter R. Russell and William Gracey.
March 1954. 9p. diagrs., photo. (NACA
RM L53L23a) IDeclassified from Confidential,
8 17, 55)

Wind-tunnel tests of the variation of total-pressure
error with angle of attack of a shielded total-
pressure tube over an angle-of-attack range of 00 to
600 and at a Mach number of 1.61 are presented.
The tests showed that the total-pressure error was
zero up to an angle of attack of 340, increased to
0.035qc, (where qc' is the indicated impact
pressure) at 480. then began to decrease and
reached a value of zero at 550 and -0. 07q,' at
59.50.


NACA RM L54C18

DYNAMIC STABILITY AND CONTROL CHARACTER-
ISTICS OF A DUCTED-FAN MODEL [N HOVERING
FLIGHT. Robert H. Kirby. April 1954. 16p.
diagrs., photos. (NACA RM L54C18) (Declassified
from Cornidential, 8. 17 '55)

This paper presents the results of an experimental
investigation of the dynamic stability and control of a
simple ducted-fan model in hovering flight and is in-
tended to provide some basic information on the sta-
bility and control of jet vertically rising airplanes in
hovering flight. The model consisted of an 18-inch-
diameter dual-rotating propeller in a shroud 4 feet
long. Control vas provided by all-movable surfaces
at the rear of the shroud. The investigation con-
sisted mainly of flight tests with the model hovering
at altitude and near the ground.


NACA Langley Field, Va.




UNIVERSITY OF FLORIDA


3 1262 08153 242 5




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