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:
AA00009235:00058

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National Advisory Committee For Aeronautics


Research A
NO.93


CURRENT NACA REPORTS

NACA RM E55BlI

FULL-SCALE PERFORMANCE STUDY OF A PRO-
TOTYPE CRASH-FIRE PROTECTION SYSTEM FOR
RECIPROCATING-ENGINE-POWERED AIRPLANES.
Dugald 0. Black and Jacob C Moser. November
i1955. 36p. diagrs., photos. INACA RM E55811I

An airplane was experimentally crashed to study the
performance of a prototype crash-lire mertng sys-
...tem for reciprocating-engine-powered airplanes.
"The results of previous experimental crashes mdi-
cate that the crash conditions imposed almost always
':: result in fire. The merting system was therefore
exposed to conditions that would adequately test its
Ability to inert and de-energize the various ignition
Sources known to cause crash fires. The fact that
fire did not occur during this crash indicated that the
.cirash-fire inerting system functioned satisfactorily
as a complete unit. The prototype inertng system
:. functioned with a rapidity equal to or greater than
S :hat of the experimental system used in the NACA
crash-fire studies.


NACA RM E55H 11

A SURVEY OF UNCLASSIFIED AXIAL-FLOW-
COMPRESSOR LITERATURE. Howard Z. Herzig
and Arthur G. Hansen. November 1955. 1, 88p
(NACA RM E55Hl1l)

A survey of unclassified axial-flow-compressor
literature is presented in the form of brief reviews
of the methods, results, and conclusions of selected
reports. The reports are organized into several
main categories with subdivisions, and frequent rel-
e. ences are made within the individual reviews to
pertinent material elsewhere in the survey.


NACA RM E55I27a

AVERAGE BOND ENERGIES BETWEEN BORON
AND ELEMENTS OF THE FOURTH, FIFTH, SIXTH,
AND SEVENTH GROUPS OF THE PERIODIC TABLE.
Aubrey P. Altshuller. November 1955. 7p. tab.
(NACA RM E55I27a)

The average bond energies Dgm(B-Z) for boron-
containing molecules have been calculated by the
Pauling geometric-mean equation. These calculated
bond energies are compared with the average bond
energies Dexp(B-Z) obtained from experimental
data. The higher values of Dexp(B-Z) in compar-
ison with Dm(B-Z) when Z is an element in the
fifth, sixth, or seventh periodic group may be attrib-
uted to resonance stabilization or double-bond char-
Sacter.


bstracts /
NOVEMBER 30 ~ 55

NACA TM 1384 DEC 5 955

METALLOGRAPHY OF ALUMINUM AN Jb
ALLOYS. USE OF ELECTROLYTIC POL H
IMetallographie de I'aluminium et de ses at 'ges.
Emploi du polissage blectrolytique). P. A. Ja et.
November 1955. ii, 80p. diagrs., photos., tabs.
(NACA TM 1384. Trans. from Office National
d'Etudes et de Recherches AEronautiques, Pub.51,
1952)

Recent methods are described for electropolishing
aluminum and aluminum alloys. Numerous refer-
ences are included of electrolytic micrographic
investigation carried out during the period 1948 to
1952. A detailed descrilM n-onf-a ep7im.t --rT-'--: -
trolytic polishing unit, s table drPmchcr6graliic'eix-
amination of aluminum a d its'aJl6ys, is included.



I --.---- --
NACA TN 3413 U.. DEPOtTORY

INVESTIGATION OF THE USE OF A RUBBER ANA-
LOG IN THE STUDY OF STRESS DISTRIBUTION IN
RIVETED AND CEMENTED JOINTS. Louis R.
Demarkles, Massachusetts Institute of Technology.
November 1955. 97p. diagrs., tabs.
(NACA TN 3413)


Results are presented of an investigation made to
study the stress distribution within cemented and
riveted joints by use of an analogous joint con-
structed of a highly flexible material. Displacement
measurements obtained from foam-rubber analogs,
and rational though not rigorously sound formulas
for shear stress distribution in joints, are given.




NACA TN 3469

SUMMARY OF RESULTS OBTAINED BY
TRANSONIC-BUMP METHOD ON EFFECTS OF
PLAN FORM AND THICKNESS ON LIFT AND DRAG
CHARACTERISTICS OF WINGS AT TRANSONIC
SPEEDS. Edward C. Polhamus. November 1955.
33p. diagrs., tab. (NACA TN 3469. Supersedes
RM L51H30)

This paper presents a summary of the effects of
plan form and thickness on the lift and drag charac-
teristics of wings at transonic speeds and compari-
sons with subsonic, transonic, and supersonic
theories. The data considered in this summary were
obtained during a transonic research program con-
ducted in the Langley high-speed 7- by 10-foot tunnel
by the transonic-bump method. The Reynolds num-
bers of the tests were generally less than 1 x 106.


.AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW,
THE REPORT TITLE AND AUTHOR.
atL. io -


WASHINGTON 25, D. C.. CITING CODE NUMBER ABOVE EACH TITLE;





2


NACA TN 3492

DETERMINATION OF INFLOW DISTRIBUTIONS
FROM EXPERIMENTAL AERODYNAMIC LOADING
AND BLADE-MOTION DATA ON A MODEL HELI-
COPTER ROTOR IN HOVERING AND FORWARD
FLIGHT. Gaetano Falabella, Jr., and John R.
Meyer, Jr., Massachusetts Institute of Technology.
November 1955. 184p. diagrs., photos., tab.
(NACA TN 3492)

Inflow distributions, azimuth and spanwise, were
determined analytically from measured pressure
distributions and blade-motion data on a model heli-
copter rotor blade under hovering and simulated
forward-flight conditions. Pressures and corre-
sponding blade flapping were recorded for various
rotor conditions at tip-speed ratios of 0.10 to 1.00.
Supplementary information concerning reverse-flow
effects on offset-blade motion, measured forces and
moments on a typical offset model rotor, and addi-
tional recorded pressure data are also included.








NACA TN 3524

THE EFFECT OF REYNOLDS NUMBER ON THE
STALLING CHARACTERISTICS AND PRESSURE
DISTRIBUTIONS OF FOUR MODERATELY THIN
AIRFOIL SECTIONS. George B. McCullough.
November 1955. 24p. diagrs., tabs.
(NACA TN 3524)

Low-speed measurements of the lift, drag, pitching
moment, and pressure distribution of the NACA
0008, 0007.5, 0007, and 0006 airfoil sections are
presented for Reynolds numbers from 1.5 to 6 mil-
lion. It is shown that the flow over these sections
underwent a change at some value of the lift coeffi-
cient which depended on the airfoil thickness ratio
and Reynolds number. The effect of the flow change
on maximum lift was small.







NACA TN 3525

VORTEX INTERFERENCE ON SLENDER AIR-
PLANES. Alvin H. Sacks. November 1955. 19p.
diagr. (NACA TN 3525)

Formulas are developed for the forces and moments
due to vortex interference on a slender wing-body-
tail combination of general cross section performing
quasi-stationary maneuvers. It is found that in
steady straight flight the interference lift depends
only on the impulse of each shed vortex and its
image vortex in a transformed circle plane, this
quantity to be determined at the wing trailing edge
and at the base of the configuration.


NACA
RESEARCH ABSTRACTS NO.93


NACA TN 3526

FLIGHT CALIBRATION OF FOUR AIRSPEED SYS-
TEMS ON A SWEPT-WING AIRPLANE AT MACH
NUMBERS UPTO 1.04 BY THE NACA RADAR-
PHOTOTHEODOLITE METHOD. Jim Rogers
Thompson, Richard S. Bray, and George E. Cooper.
November 1955. 41p. diagrs., photos.. tab.
(NACA TN 3526. Supersedes RM A50H241

The characteristics of four different airspeed sys-
tems installed in a swept-wing airplane have been
investigated in flight up to 1.04 Mach number by the
NACA radar-phototheodolite method of airspeed
calibration. The variations of static-pressure de-
fect per unit indicated impact pressure with Mach
number and a limited amount of information on the
effect of airplane normal-force coefficient are pre-
sented for each system. The results are compared
with available theory and wind-tunnel tests of the
isolated heads.






NACA TN 3547

AERODYNAMIC CHARACTERISTICS OF A SMALL-
SCALE SHROUDED PROPELLER AT ANGLES OF
ATTACK FROM 00 TO 900. Lysle P. Parlett.
November 1955. 12p. diagrs. (NACA TN 3547)

Tests have been performed to determine the effects
of airspeed and angle of attack on the lilt, drag, and
pitching moment of a shrouded-propeller model,
having a shroud length of about two-thirds of the
propeller diameter, over an angle-of-attack range
from 0 to 900. Tests were made of the complete
model with the propeller operating and also of the
shroud alone with the propeller removed. The effect
of inlet-lip cross-sectional radius on the static-
thrust characteristics was also studied.






NACA TN 3548

FLIGHT INVESTIGATION AT MACH NUMBERS
FROM 0.6 TO 1.7 TO DETERMINE DRAG AND
BASE PRESSURES ON A BLUNT-TRAILING-EDGE
AIRFOIL AND DRAG OF DIAMOND AND CIRCULAR-
ARC AIRFOILS AT ZERO LIFT. John D. Morrow
and Ellis Katz. November 1955. 19p. diagrs.,
photos. (NACA TN 3548. Supersedes RM L50E19a)

Results of an exploratory free-flight investigation at
zero lift of several rocket-powered drag-research
models having rectangular 6-percent-thick wings are
presented for a Mach number range of 0.6 to 1.7.
Wings having diamond, circular-arc, and blunt-
trailing-edge airfoil sections were tested. Pres-
sures over the base of the blunt-trailing-edge airfoil
were measured. The drags of all the models were
measured and are compared with theory in this
paper.





NACA
RESEARCH ABSTRACTS NO 93

NACA TN 3569

COMPRESSIBLE LAMINAR BOUNDARY LAYER
AND HEAT TRANSFER FOR UNSTEADY MOTIONS
OF A FLAT PLATE. Simon Ostrach. November
1955. 26p. diagrs., tab. INACA TN 3569)

The laminar compressible boundary layer and heat
transfer over an isothermal semi-inlmite flat plate
moving with a time-dependent velocity has neen ana-
lyzed. First-order deviations from the quasi-
steady velocity and temperature profiles and
boundary-layer characteristics have been computed.
A plate oscillating about a steady velocity is consid-
ered as an example.



NACA TN 3574

ACOUSTIC ANALYSIS OF RAM-JET BUZZ. Harold
Mirels. November 1955 33p diagrs
(NACA TN 3574)

A one-dimensional analysis of ram-jet buzz is pre-
sented. It is assumed that the buzz has a linear
instability origin and that the combustion chamber is
of constant area. The configuration is shown to be
unstable when the real part ol the acoustic impedance
of the inlet is greater than a term of the order of the
combustion-chamber Mach number. Computations
indicate that burning with a fixed planar flame front
and constant heat release per unit mass increases
the stable operating range.



NACA TN 3583

CHARTS OF BOUNDARY-LAYER MASS FLOW AND
MOMENTUM FOR INLET PERFORMANCE ANALY-
SIS MACH NUMBER RANGE, 0.2 TO 5.0 Paul C.
Simon and Kenneth L Kowalski. November 1955.
32p. diagrs., tab. (NACA TN 3583)

Significant flow parameters for various fractions of
a turbulent boundary layer are presented in chart
form for a number of power-law velocity profiles
and a range of Mach numbers up to 5.0. Estimates
of auxiliary inlet mass flow or momentum may easily
be made. Application of the charts to inlets of
arbitrary shape and to the determination of the pres-
sure recovery of rectangular normal-shock inlets
immersed in boundary layer is described



NACA TN 3586

IMPINGEMENT OF WATER DROPLETS ON NACA
65A004 AIRFOIL AT 00 ANGLE OF ATTACK.
Rinaldo J Brun and Dorothea E Vogt. November
1955. 28p. diagrs. INACA TN 35861

The trajectories of droplets in the air flowing past
an NACA 65A004 airfoil at an angle of attack of 00
were determined. The amount of water in droplet
form impinging on the airfoil, the area of droplet
impingement, and the rate of droplet impingement
per unit area on the airfoil surface were calculated
from the trajectories and presented to cover a large
range of flight and atmospheric conditions. These
impingement characteristics are compared briefly
with those previously reported for the same airfoil
at angles of attack of 40 and 8.


3

BRITISH REPORTS




N-40040'

National Gas Turbine Establishment iGt. Brit.)
AN EXPERIMENTAL INTRODUCTION TO THE JET
FLAP. N. A. Dimmock. July 1955. 68p. diagrs.,
photos., tabs. INGTE R. 175)

Results are given of two airfoils, each having 12.5-
percent-thick elliptical cross section with a narrow
full-span jet slot at the trailing edge, the jet deflec-
tions being, respectively. 900 and 31.40. The val-
ues of the force and moment coefficients and deriv-
atives agree with those suggested by theory in a pre-
vious report. Support is given to the thrust hypoth-
esis in that the measured thrust was greater, under
appropriate conditions, than the reaction component
from the deflected jet. The losses in the system
are considered and some are mvestigated, those due
to Reynolds number and jet entrapment effects being
included. Influence of ground on lift and center of
lift was measured and found not to be prohibitive.



N -40048'

Royal Aircraft Establishment (Gt Brit.)
TESTS OF HUMIDITY EFFECTS ON FLOW IN A
WIND TUNNEL AT MACH NUMBERS BETWEEN
2.48 AND 4. R. J. Monaghan. January 1955. 33p.
diagrs. (RAE Tech. Note Aero 2358)

Static and pilot pressure distributions were meas-
ured in the working section of a 5-in. by 5-in.
supersonic wind tunnel at nominal Mach numbers of
2.48, 3.25, and 4, over ranges of absolute humidity
at the inlet from 5 x 10-5 to 3 x 10-3. Previous
work indicates that a condensation shock would occur
in the nozzle. For a stagnation pressure of I at-
mosphere and stagnation temperatures giving zero
heat-transfer conditions at the walls, no humidity
effects were discernible if the absolute humidity was
less than 2 x 10-4 at M = 2.48, 3 x 10-4 at M = 3.12,
and about 5 x 10-4 at M = 3.8.



N-40068

Aeronautical Research Council (Gt Brit. I
TESTS ON A SWEPT-BACK WING AND BODY WITH
ENDPLATES AND WING TIP TANKS IN THE COM-
PRESSED AIR TUNNEL. C. Salter and R. Jones.
APPENDIX COMPARISON BETWEEN THE MEAS-
URED LIFT AND DRAG AND CALCULATED VAL-
UES FOR THE WING WITH TIP TANKS. J. Weber.
1954. 26p. diagrs.. tabs. IARC CP 196)

Results are given of experiments to determine the
effect on lift. drag and pitching moment, of wing lip
tanks and of two sizes of end plates on a tapered
swept wing model. The tests were undertaken pri-
marily to extend the range of Reynolds number for
checks on previous theoretical work. As regards
lilt and pitching moment, the effects are found to be
fairly well defined. The drag characteristics are
less consistent, but it seems that end plates have
the effect of reducing the drag of the model over
quite a large range of CL. This does not apply in
the case of the wing tip tanks.






4



N-40069*

Aeronautical Research Council (Gt. Brit.)
WIDE RANGE AMPLIFIER FOR TURBULENCE
MEASUREMENTS WITH ADJUSTABLE UPPER
FREQUENCY LIMIT. H. Schuh and D. Walker.
1955. 42p. diagrs. (ARC CP 198)

Requirements are discussed for an amplifier suit-
able for subsonic and supersonic turbulence work
with hot wires. An amplifier is described which has
a frequency range from 1.4 c/s to 50 kc/s, dealing
with a range of thermal time lag from 0.1 m. s. to
5 m. s. An iron dust-cored inductance is used to
give the required compensation for thermal lag, the
circuit being a modification of Dryden's circuit.
The upper frequency cutoff is adjustable in six steps
from 1.5 kc/s to 50 kc/s. The output can be applied
to a thermocouple meter and to an oscilloscope.




N-40080*

Aeronautical Research Council (Gt. Brit.)
FLIGHT TESTS AT TRANSONIC SPEEDS ON FREE-
LY FALLING MODELS. PARTS I TO V. Edited by
C. Kell. PART I HISTORICAL. C. Kell. PART
I EQUIPMENT AND TECHNIQUE. C. Kell and
J. Swan. PART III -DRAG EXPERIMENTS.
C. Kell and F. Smith. PART IV FLUTTER EX-
PERIMENTS. W. G. Molyneux and E. W. Chapple.
PART V NOTES ON THE ACCURACY OF THE
FREELY FALLING MODEL EXPERIMENT.
T. F. C. Lawrence. 1955. 32p. diagrs., photos.,
tab. (ARC R & M 2902. Supersedes RAE Tech.
Memo. Aero 308)

Basic bodies carrying the airfoils to be tested were
released from an aircraft flying at height, and accel-
erated under the influence of gravity through the
transonic speed range. Radar recorded the flight
path and telemetering equipment carried within the
body transmitted information to a ground station
during the free fall. This work started in 1943 and
was brought to a close in 1949.




N-40082*

Aeronautical Research Council (Gt. Brit.)
OBSERVATIONS OF THE FLOW ROUND A TWO-
DIMENSIONAL AEROFOIL OSCILLATING IN A
HIGH-SPEED AIR STREAM. A. Chinneck, D. W.
Holder, and C. J. Berry. 1955. 18p. diagrs.,
photos., tab. (ARC R &M 2931. Supersedes ARC
15, 141; FM 1779 & 0.1006)

Photographs have been taken of the flow around a
10-percent-thick RAE 104 airfoil performing pitch-
ing oscillations at low values of the frequency pa-
rameter in subsonic and supersonic airstreams.
Apart from a difference of phase, the general flow
pattern appeared to be similar to those observed r,
steady motion, the pattern for a particular instanta-
neous incidence of the oscillation resembling that
for steady motion at a different incidence. It is
suggested that, for the range of frequency parameter
covered, the observed phase lag of the flow pattern
corresponds to the lag in the circulation.


NACA
RESEARCH ABSTRACTS NO.93


N-40084*

Aeronautical Research Counc il IGt. Brit.I
THE AERODYNAMIC EFFECTS OF ASPECT RATIO
ON FLUTTER OF UNSWEPT WINGS. W. G.
Molyneux and E. W. Chapple. 1955. 12p. diagrs..
tab. (ARC R & M 2942; 15.609. Supersedes RAE
Structures 135)

A method is described for the direct measurement of
the aerodynamic effects of aspect ratio on wing flut-
ter. The method requires the use of stilf (virtually
rigid) wings flexibly mounted at the root. Details
are given of tests on untapered. aun-wept wings with
freedoms in modes of linear Ilexure and uniform
pitch. A comparison is made between measured
values of the flutter characteritlcs and the values
calculated using an aerodynamic theory for oscil-
lating wings of finite aspect ratio, and reasonable
agreement for flutter speeds andr frequencies is
obtained.


N-40186*

Royal Aircraft Establishment IGt. Brit.I
MEASUREMENTS OF PITCHING MOMENT DERIVA-
TIVES FOR A SERIES OF RECTANGULAR WINGS
AT LOW WIND SPEEDS. P. R. Guyett and D. E. G.
Poulter. June 1955. 52p. di3grs., tabs. (RAE
Structures 185)

The direct aerodynamic moments for pitching oscil-
lations have been measured on a series of rectangu-
lar wings having aspect ratios between 2 and 8 for
axis positions at the wing leading edges and trailing
edges. Two of the wings were also tested with
single end plates which were aerodynamically effec-
tive in doubling the wing geometric aspect ratio.
The measurements were made at low speeds in an
open jet wind tunnel and covered the range of fre-
quency parameter (based on wing chord) 0.13 to 0.39.
The results are in general arc-ement with theoreti-
cal results due to Lawrence and Gerber. Similar
tests were also made on a wing fitted qwth two end
plates in an attempt to obtain results for two-
dimensional flow. The results do not agree with
other experimental results and two-dimensional
theoretical values and indicate that wind-tunnel in-
terference is important for this test configuration.


N-40187'

Royal Aircraft Establishment iGt. Brit.)
THE MECHANICAL PROPERTIES AND STRUCTURE
OF CONTINUOUSLY CAST A.C.9 ALUMINIUM
ALLOY TUBES. P. C. Bradley and D. Bunting.
June 1955. 10p. diagr., photos.. tabs. IRAE Tech.
Note RPD 1221

The mechanical properties of a large diameter tube
continuously cast in aluminum alloy A.C. 9 have been
investigated at room and at elevated temperature, in
connection with the use of such tubes as fuel or
oxidant tanks for rocket motors. The variation in
structure of the alloy across a section is described
and discussed with reference to the possible detri-
mental effect of primary silicon crystals on the
mechanical propertle;. The operating temperature
should nol be allowed to exceed about 1800 C if tanks
of thi material are to be used more than once.





NACA
RESEARCH ABSTRACTS NO 93


N-40188'

Royal Aircraft Estaulihnhment IGt. Brit)
THE EFFECTS OF TAPER ON THE SUPERVELOC-
ITIES ON THREE-DIMENSIONAL WINGS AT ZERO
INCIDENCE. K. W. Newby. June 1955. 120p.
diagrs.. Lab. (RAE Aero 25441

Relationships ha\e been derived for expressing the
velocities on three-diniensioural tapered wings at
zero incidence in terms o the velocities on un-
tapered infinite swept wings. The theoretical in\es-
tigation of the effects of taper is confined to simple
wings having airluil sections formed by cubic or
parabolic arce; some experimental evidence is given
to show that the results of this investigation can
probably be applied quantitatively to wings having
conventional airfoil sections. The results given in
this report show that plan form and thickness taper
have a marked effect on the velocities near the cen-
ter of a wing, but that these effects decrease with in-
crease of s\teepback. A calculation method is out-
lined in section 4.26 of the text for applying the re-
sults obtained for wings hat ing parabolic arc airfoil
sections, to wings having arbitrary section shapes.


N-40189'

Royal Aircralt Estclishment IGt. Brit.)
RECORDING AND PROCESSING FLIGHT TEST
DATA BY DIGITAL METHODS. E. J. Petherick.
(Prepared for AGARD Flight Test Panel). April
1955. 12p. photos. iRAE Tech. Note MS 20)

This note reviews current developments in the re-
cording and processing of flight test data by digital
methods. It first describes four assessors which
facilitate conversion of strip chart and kinetheodolite
records to typewritten or punched card form. It
then details some coded scales which can be read
automatically, and finally it describes the incorpora-
tion and use of such scales in a digital recording
system intended for airborne use



N-40190'

Royal Aircraft Establishment (Gt. Brit.)
WIND TUNNEL TESTS ON A 6 FT DIAMETER HEL-
ICOPTER ROTOR. T. B. Owen, R. A. Fail, and
R. C. W. Eyre. May 1955. 33p. diagrs., tabs.
(RAE Tech. Note Aero 2378)

Thrust, torque, and flapping angle have been meas-
ured on a 6-foot diameter rotor over a range of
blade angle, shalt inclination, and lip speed ratio for
comparison with the 12-foot diameter rotor previ-
ously tested in the 24-foot tunnel. In addition to
tests in the 24-foot tunnel, the 6-foot diameter rotor
was also tested in the No. 2 11-1 2 foot tunnel to
investigate tunnel constraint. Brief investigations
were made of support interference and blade twist-
ing. There are small discrepancies both as regards
tunnel corrections and as regards the comparison of
the 6-foot and 12-foot diameter rotors in the un-
stalled operating range. Possible reasons are dis-
cussed but small unexplained discrepancies remain.
Blade stalling has larger effects on the 6-foot diam-
eter rotor but owing to the progressive nature of the
phenomenon it is not possible to define any precise
limits to the ranges of validity of the results on the
two rotors.


5


N-40191'

Aeroplane and Armament Experimental Establish-
ment (Gt. Brit.) AN EXPERIMENTAL INVESTIGA-
TION INTO THE PERFORMANCE OF A HELICOP-
TER FOLLOWING SUDDEN REDUCTION IN POWER.
G. W. Langdon. August 4, 1955. 12p. diagrs.
(AAEE, Res 289).

The performance of a single rotor helicopter fitted
with'a throttle override has been measured under
conditions simulating the failure of one engine of a
multiengined helicopter during take-off. Records of
the motion are included and the experimental results
are compared with theoretical predictions.








N-40192'

Royal Aircraft Establishment (Gt. Brit.)
INVESTIGATION OF THE FATIGUE OF EXTRUDED
TUBULAR BOOMS. W. A. P. Fisher and
H. Yeomans. June 1955. 15p. diagrs., photos.,
tabs. (RAE Tech. Note Structures 162).

The presence of the unmachined extruded surface of
the bore in the tubes, as used for the 'Viking" and
"Valetta" spar booms, has a marked adverse effect
on the basic fatigue strength of the tube. The fail-
ures of the necked specimens show that the flaws at
the inner surface are a source of fatigue. Scatter in
the endurance of tubular spar booms is probably
largely due to the chances of such flaws occurring at
the side of a transverse hole. The specimens were
made from aluminum alloy extruded tube DTD 364.








N-40193'

Marine Aircraft Experimental Establishment. (Gt.
Brit.) INVESTIGATION OF HIGH LENGTH BEAM
RATIO SEAPLANE HULLS WITH HIGH BEAM
LOADINGS HYDRODYNAMIC STABILITY.
PART 17 THE STABILITY AND SPRAY CHARAC-
TERISTICS OF MODEL M. J. K. Friswell, D. M.
Ridland, and A. G. Kurn. April 1955. 24p. diagrs.,
photos., tabs. (MAEE F. Res 253)

In this report results are presented of limited tests
on the hydrodynamic characteristics of model M of
the series, these tests being designed solely to pro-
vide information on the interactions of the different
relevant parameters. The model has a length-to-
beani ratio of 13 (the forebody being 6 beams in
length and the afterbody 7 beams), no 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 load-
ing. A short discussion of the results is also
included.






6



N-40197*

Royal Aircraft Establishment (Gt. Brit.)
CORRELATED FATIGUE DATA FOR AIRCRAFT
STRUCTURAL JOINTS. R. B. Heywood. June 1955.
16p. diagrs., tab. (RAE Structures 184)

Results of fatigue tests carried out at RAE on typi-
cal aircraft wing structural joints are correlated to
give an indication of general fatigue behavior. The
results are plotted in the form of S Log N curves,
and these indicate that the mode of behavior cannot
be attributed to any single factor, such as the type of
aluminum alloy, the ultimate tensile strength, or the
mean stress of the fatigue cycle. The detailed meth-
od of design undoubtedly has a predominant influence
on behavior, but this quality is not revealed by a
broad classification according to the proportion of
load transmitted at holes.




N-40198

Royal Aircraft Establishment (Gt. Brit.)
A TECHNIQUE FOR THE MEASUREMENT OF
PRESSURE DISTRIBUTION ON OSCILLATING
AEROFOILS, WITH RESULTS FOR A RECTANGU-
LAR WING OF ASPECT RATIO 3.3. W. G. Molyneux
and F. Ruddlesden. June 1955. 24p. diagrs., tabs.
(RAE Tech. Note Structures 164)

Details are given of a strain-gage pressure trans-
ducer that has been developed for measurements of
pressure distribution on oscillating airfoils in low
speed wind tunnels. The transducer characteristics
are shown to be well suited to oscillatory measure-
ments, and in particular the transducer output can be
measured directly on a sensitive galvanometer with-
out the need for preamplification. As an illustration
of the use of the transducer, pressure measurements
have been made in the RAE 5-foot diameter open jet
wind tunnel on a rectangular wing of aspect ratio 3.3
oscillating in modes of pitch and roll. Values for
the aerodynamic derivatives have been obtained from
the integrated pressure distributions, and are com-
pared with those derived from overall force meas-
urements and with theoretical values. The measured
values are in close agreement but there are some
discrepancies with theory that are thought to be due
to a wind-tunnel interference effect.





N-40199*

Royal Aircraft Establishment (Gt. Brit.)
THE INFLUENCE OF PRE-LOADING ON THE FA-
TIGUE LIFE OF AIRCRAFT COMPONENTS AND
STRUCTURES. R. B. Heywood. June 1955. 27p.
diagrs., photos., tabs. (RAE Structures 182)

Tests on aircraft components and structures are
described which show that preloading can have a


NACA
RESEARCH ABSTRACTS NO. 93



marked influence on fatigue behavior. Tensile pre-
loading may increase the life in one instance a
hundredfold improvement was obtained and com-
pressive preloading may reduce the life. The effect
is attributed to residual stresses and to load re-
distributions induced by preloading.



N-40205*

Ministry of Supply (Gt. Brit.)
AN ABSORPTIOMETRIC METHOD FOR THE DE-
TERMINATION OF CHROMIUM IN
METHACRYLATO-CHROMIUM TREATED GLASS
FABRIC. E. I. McLauchlan. August 1955. 4p.
diagr. (MOS AID Chem. 6)

An absorptiometric method for the determination of
chromium in methacrylato-chromium treated glass
fabric is described. A wet-oxidation attack is uti-
lized to remove the chromium from the glass fibers
and to oxidize the tervalent chromium to sexavalent
chromium for subsequent absorptiometric deter-
mination as the diphenylcarbazide complex.





N-40207*

Royal Aircraft Establishment (Gt. Bnt.)
OPTIMUM DESIGNS FOR REINFORCED CIRCULAR
HOLES. E. H. Mansfield. June 1955. 26p. diagrs.
(RAE Structures 183)

The design of reinforced circular holes in an infinite
sheet is considered theoretically. The stress sys-
tem in the main body of the sheet is assumed to be
one in which the principal stresses are in the ratio
1:-1 (that is, shear), 1:0 (that is, tension), 1:1 or
1:1/2. The reinforcement may vary round the hole
and families of such reinforcements with constant
total weight are considered; the peak stresses in the
sheet are evaluated so that optimum weight-strength
designs are determined.




N-40209*

Royal Aircraft Establishment (Gt. Brit.)
STATIC ELECTRICITY IN AIRCRAFT FUEL TANKS.
F. L. Holmes and D. T. Sharwood. May 1955. 10p.
diagrs., tab. (RAE Tech. Note EL.82; Tech. Note
Mech. Eng. 201)

Tests were made to measure the static field strength
generated in aircraft fuel tanks by foaming of the
fuel. A type of apparatus for measuring static field
strength is described with details of its application
to these tests. The results show that the field
strength is less than 1 volt per centimeter which is
many times less than that required for flashover and
the ignition of a fuel vapor air mixture.





NACA
RESEARCH ABSTRACTS NO.93



UNPUBLISHED PAPERS




N-22024'

FLOW NEAR A HEATED SOLID BODY IN A STAND-
ING ACOUSTIC WAVE. P. N. Kubanskii. October
1955. lip. diagrs. (Trans. of Zhurnal Tekhnich-
eskoi Fiziki, v. 22, no.4, April 1952, p.585-592)

Results of experiments carried out in standing waves
generated by vibrations of finite amplitude are pre-
sented. Two types of waves are considered:
(1) forced standing waves, formed when the frequen-
cy of outside vibrations does not correspond with the
natural frequency of vibration of the radiating sys-
tem, and (2) standing waves which form as the re-
sult of coincidence of frequency of outside vibration
with natural frequency of vibration of the radiating
system. Optical methods are used in the examina-
tion of the flow and the presence of higher harmonics
in the standing acoustical wave is observed. The
deformation of the waves is also studied.



N-34452'

SOFT ROT. DESTRUCTION OF WOOD THROUGH
COMMON FUNGI. (Moderfaule. Die Zersetzung
von Holz durch niedere Pilze). W. P. K. Findlay and
J. G. Savory. October 1955. 12p. photos., tab.
(Trans. from Holz als Roh- und Werkstoif, v.12,
no.8, August 1954, p.293-296).

II wood is exposed to humid weather for some length
of time, a softening of the surface takes place which
is due to a fungus attack. This phenomenon is of
minor significance except in the case of cooling
towers and similar industrial water-cooling plants.
Soft rot, as this fungus is generally called, was pro-
duced according to a method by Abrams. The fungi
causing soft rot are reviewed and preservatives for
protection of wood are discussed, but further inves-
tigations are required in order to indicate effective
methods of protection of wood from soft rot.



N-39717'

EFFECT OF ACOUSTICAL VIBRATIONS OF FINITE
AMPLITUDE ON THE BOUNDARY LAYER. P. N.
Kubanskii. October 1955. lip. diagrs. (Trans. of
Zhurnal Tekhnicheskoi Fiziki, v.22, no.4, April 1952,
p.593-601)

It is concluded that acoustical standing waves of
finite amplitude can produce currents near the wall
of a solid body, even when the body is located in a
stream. A considerable intensity of acoustical
vibration is necessary in order to produce acoustical
flow near the walls of a solid body in a stream. The
acoustical streams produced near the walls of solid
bodies exert an influence on the boundary layer
which surrounds the body. There is a possibility of
using acoustical vibrations to control phenomena
which occur in the boundary layer. Such phenomena
could be directed toward the desired side of a physi-
cal body, depending upon the condition of the bound-
ary layer.


7



DECLASSIFIED NACA REPORTS




NACA RM A50J26a

AERODYNAMIC CHARACTERISTICS INCLUDING
PRESSURE DISTRIBUTIONS OF A FUSELAGE AND
THREE COMBINATIONS OF THE FUSELAGE WITH
SWEPT-BACK WINGS AT HIGH SUBSONIC SPEEDS.
Fred B. Sutton and Andrew Martin. February 6,
1951 117p. diagrs., photos., tabs.
(NACA RM A50J26a)
(Declassified from Confidential, 10.,14/55)

As part of an NACA transonic research program,
three sweptback wings with a fuselage were inves-
tigated over a Mach number range from 0.40 to 0.94.
These model wings had NACA 65A006 sections par-
allel to the plane of symmetry. One of the model
wings was swept back 350 and had an aspect ratio of
6; the other two were swept back 450 and had aspect
ratios of 4 and 6. Force and pitching-moment data,
tabulated pressure measurements, downwash and
dynamic-pressure characteristics, and tuft studies
are presented. The approximate effects of wing
elasticity on liut and moment data are also shown.





THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL 11/14/55,
AND ARE UNAVAILABLE:

RM L8K18a
RM L52K24a








THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL, 11/14/55







NACA RM A7K28

HIGH-SPEED STABILITY AND CONTROL CHARAC-
TERISTICS OF A FIGHTER AIRPLANE MODEL
WITH A SWEPT-BACK WING AND TAIL. Charles
P. Morrill, Jr., and Lee E. Boddy. April 14, 1948.
47p. diagrs., photos. (NACA RM A7K28)

Wind-tunnel tests were made at high subsonic Mach
numbers of a model of a pursuit airplane with a 350
sweptback wing and tail. Data are included in the
report which show the basic characteristics of the
model; the control and hinge-moment characteristics
of the horizontal tail, elevator, and aileron; the ef-
fect of a wing leading-edge slat; and the effect of a
fuselage-side dive brake.




UNIVERSITY
8 I ll I 1111111

3 1262 0
NACA RM A9K02

INVESTIGATION OF DOWNWASH AND WAKE
CHARACTERISTICS AT A MACH NUMBER OF 1.53.
III SWEPT WINGS. Edward W. Perkins and
Thomas N. Canning. February 23, 1950. 41p.
diagrs., tab. (NACA RM A9K02)

The results of an experimental investigation of the
downwash and wake characteristics behind two
highly swept wings in a supersonic stream are pre-
sented. The leading-edge sweep angles of the two
wings were 630 and 63045', the aspect ratios were
3.50 and 1.66, and the corresponding taper ratios
were 0.25 and 1.00. The tests were made at a Mach
number of 1.53 and Reynolds numbers of 1.4 million
and 2.6 million, respectively. A comparison be-
tween experimental and theoretical values of the rate
of change of downwash angle with angle of attack at
zero lift is made.






NACA RM A52K12

TECHNIQUES FOR DETERMINING THRUST IN
FLIGHT FOR AIRPLANES EQUIPPED WITH
AFTERBURNERS. L. Stewart Rolls, C. Dewey
Havill, and George R. Holden. January 1953. 27p.
diagrs., photos. (NACA RM A52K12)

An experimental technique has been developed which
enables a determination of the net thrust for an
afterburner-equipped airplane in flight. Measure-
ments from a swinging pitot-static pressure and
total temperature probe are used to determine the
gross thrust, total air-flow rate, and net thrust.
Details are also presented for an air-cooled fixed-
pressure probe for the determination of basic engine
thrust.






NACA RM A52K13

TESTS IN THE AMES 40- BY 80-FOOT WIND TUN-
NEL OF AN AIRPLANE MODEL WITH AN ASPECT
RATIO 4 TRIANGULAR WING AND AN ALL-
MOVABLE HORIZONTAL TAIL HIGH-LIFT DE-
VICES AND LATERAL CONTROLS. Ralph W.
Franks. February 1953. 45p. diagrs., photo., 2
tabs. (NACA RM A52K13)

Tests have been made of a model consisting of a tri-
angular wing in combination with a fuselage of fine-
ness ratio 12.5; a thin, triangular, vertical tail with
a constant-chord rudder; and a thin, unswept, all-
movable horizontal tail. The wing had an NACA
0005 modified section and was equipped with slotted
inboard and plain outboard flaps. Tests were made
with the wing-fuselage-vertical-tail configuration in
addition to the tests of the complete model. The
results of tests of lateral and directional controls,
the inboard flaps as a high-lift device, and the out-
board [laps as a high-lift device are presented. The
Reynolds number, based on the win; mean aerody-
namic chord, was approximately 10.9 million and
the Mach number was about 0.13.


TY OF FLORID

1111111
8153 2Z
I


A
RESEARCH ABSTRACTS NO 93

3 8
NACA RM A53C19

EXPERIMENTAL INVESTIGATION OF THE EF-
FECTS OF PLAN-FORM TAPER ON THE AERO-
DYNAMIC CHARACTERISTICS OF SYMMETRICAL
UNSWEPT WINGS OF VARYING ASPECT RATIO.
Edwin C. Allen. May 1953. 53p. diagrs photos..
tab. (NACA RMA53C19)

This report presents results of tests of a series of
symmetrical, unswept. 8-percent-thick wings of
varying aspect ratio and taper ratio The wings
were tested in combination ith four ddl'erent bodies
of revolution over a Mach number range from 0.40
to 0.94 with a corresponding Reynolds number range
from 2.58 million to 5.90 million The lift, drag,
and pitching-moment data are presented for wings
of aspect ratios 2, 3, and 4 and for taper ratios of
0.20 to 1.00.






NACA RM E51K15

COMPARISON OF LOCKED-ROTOR AND WIND-
MILLING DRAG CHARACTERISTICS OF AN AXIAL-
FLOW-COMPRESSOR TYPE TURBOJET ENGINE.
K. R. Vincent, S. C. Huntle. and H D Wilsted.
January 1952. lOp. diagrs INACA RM E51K15)

The internal drag of an axial-flow turbojet engine
with the rotor locked in place to prevent uindmilling
and with the engine windmilling was obtained over a
range of simulated Mach numbers The corrected
internal drag of the engine with the locked rotor was
210 pounds or only 46 percent of the windmilling
drag at a flight Mach number of 0.8.





NACA RM E53K06

ANALYTICAL STUDY OF LOSSES AT OFF-DESIGN
CONDITIONS FOR A FIXED-GEOMETRY TURBINE.
Warner L. Stewart and David G Evans. February
1954. 48p. diagrs., tab. (NACA RM E53K06)

An analytical investigation was made to determine
the.off-design loss characteristics of a fixed-
geometry turbine of which the experimental per-
formance was known. The method of analysis
utilized an effective loss parameter and assumed
that the velocity normal to the blade entrance angle
was lost as a total-pressure loss. The method also
assumed constant tangential component of velocity
between the station just upstream and just down-
stream of the stator and rotor trailing edge. Good
correlation between the analytically and experi-
mentally obtained performance was found over the
entire map until limiting loading was approached.
The large decrease in efficiency at low-speed high
pressure ratios and at high-speed low pressure
ratios was found in the analysis to be almost en-
tirely due to the rotor incidence and turbine exit
whirl losses. From the results of the investigation
it was concluded that for turbines designed to oper-
ate efficiently at more than one point, the design
must compromise rotor incidence angle and exit
whirl losses.





NACA
RESEARCH ABSTRACTS NO.93




NACA RM L8KI2a

HINGE-MOMENT MEASUREMENTS OF A WING
WITH LEADING-EDGE AND TRAILING-EDGE
FLAPS AT A MACH NUMBER OF 1.93. William B.
Boatright and Robert W. Rainey January 14, 1949.
12p. diagrs., tab. (NACA RM L8K12a)

Hinge-moment data for a wing with leading-edge and
trailing-edge flaps of wedge section were obtained in
the Langley 9-inch supersonic tunnel at a Mach num-
ber of 1.93 and a Reynolds number of 1.31 x 106.
Curves of hinge moment against angle of attack and
against flap deflection are shown, and the results
are compared with theory The possibility of a
linkage system to reduce control forces is discussed








NACA RM L8K17a

CONTROL EFFECTIVENESS AND HINGE-MOMENT
MEASUREMENTS AT A MACH NUMBER OF 1.9 OF
A NOSE FLAP AND TRAILING-EDGE FLAP ON A
HIGHLY TAPERED LOW-ASPECT-RATIO WING.
D. William Conner and Meade H Mitchell, Jr.
January 10, 1949. 26p. diagrs photo.
(NACA RM L8K17a)

Nose flaps and trailing-edge flaps were tested on a
low-aspect-ratio, highly tapered, half-span wing
model in the Langley 9- by 12-inch supersonic blow-
down tunnel at a Mach number of 1.9 and a Reynolds
number of 3,000,000. Lilt, drag, pitching- and
rolling-moment data for the mwng and hinge-moment
data for the flaps were obtained. All tests were
made in the presence of a fuselage.







NACA RM L8K24a

EXPERIMENTAL AND CALCULATED HINGE
MOMENTS OF TWO AILERONS ON A 42 70 SWEPT-
BACK WING AT A MACH NUMBER OF 1.9.
James C. Sivells and Kennith L. Goin. January 19,
1949. 23p. diagrs., photos. tabs.
(NACA RM L8K24a)

A 42 7 sweptback wing was tested with two types of
ailerons in the Langley 9- by 12-inch supersonic
blowdown tunnel at a Mach number of 1 9 and a
Reynolds number of 2.2 x 106. The wing had an as-
pect ratio ol 4, a taper ratio of 0.5, and an 8-
percent-thick biconvex airfoil section. The contour
of one aileron was formed by the basic airfoil con-
tour, and the other aileron had flat sides and a
trailing-edge thickness of one-half the hinge-line
thickness.


9




NACA RM L9K01a

ROCKET-POWERED FLIGHT TEST OF A ROLL-
STABILIZED SUPERSONIC MISSILE CONFIGURA-
TION. Robert A. Gardmer and Jacob Zarovsky.
January 12, 1950. 32p. diagrs., photos., tab.
(NACA RM L9K01a)

A missile research model incorporating wing-lip
ailerons and a gyro-actuated automatic roll control
was flight tested at supersonic speed. The aerody-
namic rolling derivatives for zero-lift flight were
determined from the flight record. It was concluded
that the method used in the preflight system analysis
is valid and that the gyro-actuate0 control
system provided a satisfactory ,fa g
roll stabilization in zero-lif, onic ii


B CG 5 .J5


NACA RM L9K09

AERODYNAMIC INVESTIGAT' F O A PARABOLIC
BODY OF REVOLUTION AT MA -'F MBEI OF
1 92 AND SOME EFFECTS OF AN ANNULAR JET
EXHAUSTING FROM THE BASE. Eugene S. Love.
February 8, 1950. 75p diagrs., photos., tab.
(NACA RM L9K09)

An aerodynamic investigation of a slender pointed
parabolic body of revolution was conducted at a Mach
number of 1.92 with and without the effects of an
annular supersonic jet exhausting from the base.
Measurements without the jet in operation were
made of lift, drag, pitching moment, base pres-
sures, and radial and axial pressures. With the jet
in operation, pressure measurements were made
over the rear of the body with the primary variables
being angle of attack, ratio of let velocity to stream
velocity, and ratio of pressure at jet exit to stream
pressure.





NACA RM L50K06

HORIZONTAL-TAIL EFFECTIVENESS AND DOWN-
WASH SURVEYS FOR TWO 47.70 SWEPTBACK
WING-FUSELAGE COMBINATIONS WITH ASPECT
RATIOS OF 5.1 AND 6.0 AT A REYNOLDS NUMBER
OF 6.0 x 106. Reino J. Salmi January 12, 1951
65p. diagrs., photos., 2 tabs. (NACA RM L50K06)

Results of wind-tunnel tests on two 47.70 sweptback
wing-fuselage combinations of aspect ratios 5.1 and
6.0 to determine the effects of the vertical location
of a horizontal sweptback tail on tail effectiveness
and the static longitudinal stability were presented.
The tests were made at a Reynolds number of about
6 0 x 106 (Mach number of 0.14) for various combi-
nations of leading-edge and trailing-edge flaps. The
results of airstream surveys in the region of the
tail are also presented





10





NACA RM L50K29

LOW-SPEED LONGITUDINAL AND WAKE AIR-
FLOW CHARACTERISTICS AT A REYNOLDS NUM-
BER OF 6.0 x 106 OF A 520 SWEPTBACK WING
EQUIPPED WITH VARIOUS SPANS OF LEADING-
EDGE AND TRAILING-EDGE FLAPS, A FUSELAGE,
AND A HORIZONTAL TAIL AT VARIOUS VERTICAL
POSITIONS. Roland F. Griner and Gerald V. Foster.
February 28, 1951. 66p. diagrs., photo., 3 tabs.
(NACA RM L50K29)

The results are presented of an investigation con-
ducted in the Langley 19-foot pressure tunnel at a
Reynolds number of 6.0 x 106 to determine the ef-
fects of leading-edge-flap spans on an NACA 64-
series wing swept back 520. Several of the more
satisfactory spans of leading-edge flaps were inves-
tigated with various combinations of trailing-edge
flaps, fences, a fuselage, and a horizontal tail. Sur-
veys of downwash angle, sidewash angle, and dy-
namic pressure ratio behind the wing at approxi-
mately the location of a horizontal tail are presented.







NACA RM L51I05

EFFECT OF FORMATION POSITION ON LOAD
FACTORS OBTAINED ON F2H AIRPLANES. Carl R.
Huss and Harold A..Hamer. December 1951. 15p.
diagrs., 3 tabs. (NACA RM L51105)

Plots of load factor against airplane position are
presented for three combinations of four airplanes
flying in formation. The plots show that the load-
factor trend was to increase toward the end of the
formation. Typical time histories are presented.







NACA RM L51K30

TIME HISTORIES OF MANEUVERS PERFORMED
WITH AN F-86A AIRPLANE DURING SQUADRON
OPERATIONS. Harold A. Hamer and Campbell
Henderson. February 1952. 90p. diagrs., 3 tabs.
(NACA RM L51K301

Some preliminary results of maneuvers performed
during U. S. Air Force squadron operations with an
F-86A jet-fighter airplane are presented in time-
history form. The maneuvers cover a speed range
from the stall to 530-mile-per-hour indicated air-
speed and pressure altitudes varying from sea level
to approximately 25,000 feet. Variation of the
maximum airplane linear and angular accelerations
experienced during the investigation are also pre-
sented.


NACA
RESEARCH ABSTRACTS NO. 93





NACA RM L52K07

FREE-FLIGHT INVESTIGATION AT ZERO LIFT IN
THE MACH NUMBER RANGE BETWEEN 0.7 AND
1.4 TO DETERMINE THE EFFECTIVENESS OF AN
INSET TAB AS A MEANS OF AERODYNAMICALLY
RELIEVING AILERON HINGE MOMENTS. William
M. Bland, Jr., and Edward T. Marley. January
1953. 19p. diagrs., photos. INACA RM L52K07)

An experimental investigation employing a technique
which utilized a zero-lift rocket -propelled model in
free flight has been made to determine some of the
characteristics of an inset tab as an aerodynamic
balance in the Mach number range between 0.7 and
1.4. The fixed, 0.09-chord, full-span, inset tab
that was investigated was attached to a 0.3-chord
full-span aileron on a wing of aspect ratio 3 and
taper ratio 0.6 that had the quarter-chord line swept
back 450 and NACA 65A006 airfoil sections parallel
to the model center line. Results of this investiga-
tion show that the tab was capable of balancing (trim-
ming) the aileron hinge moments throughout the Mach
number range investigated even though the effective-
ness of the tab decreased with increasing Mach num-
ber. It was shown that the aileron rolling effective-
ness was decreased considerably when the tab was
used to reduce the aileron hinge moments. The tab
when considered as a servotab was an effective aero-
dynamic balance for Mach numbers less than 1.1.
At no time during the investigation did the mass-
balanced aileron show any evidence of buzz or flutter.
I was also shown that the tab effectiveness could be
estimated with reasonable accuracy from experLmen-
tal data and from thin-airfoil theory.











NACA RM L52K25

INVESTIGATION OF THE EFFECT OF CHORDWISE
POSITIONING AND SHAPE OF AN UNDERWING NA-
CELLE ON THE HIGH-SPEED AERODYNAMIC
CHARACTERISTICS OF A 450 SWEPTBACK
TAPERED-IN-THICKNESS-RATIO WING OF AS-
PECT RATIO 6. H. Norman Silvers and Thomas J.
King, Jr. January 1953. 50p. diagrs.
(NACA RM L52K25)

An investigation at high speeds of chordwise posi-
tioning of underwing nacelles at a spanwise location
of 0.46 semispan with an ogive-cylinder shape
(fineness ratio = 9.34), an NACA 65A-series airfoil
of revolution shape (fineness ratio = 10.68), and a
modified NACA 0-series airfoil of revolution shape
reversed in direction (fineness ratio = 10.04) was
made on a small-size 450 sweptback Lapered-in-
thickness-ratio wing of aspect ratio 6.





NACA
RESEARCH ABSTRACTS NO. 93




NACA RM L53K16

AN AIR-FLOW-DIRECTION PICKUP SUITABLE FOR
TELEMETERING USE ON PILOTLESS AIRCRAFT.
Wallace L. Ikard. March 1954. 25p. diagrs.
photos. INACA RM L53K16)

A vane-type air-llow-direction pickup is described
which is suitable for telemetering angle-of-attack
and angle-of-sideslip data from rocket-propelled
pilotless aircraft models. Test results which are
presented show that the device performs well under
high accelerations and is stable throughout a Mach
number range from subsonic to above a Mach num-
ber of 2.5.





NACA RM L53K18

EXPERIMENTAL INVESTIGATION OF THE OSCLL-
LATING FORCES AND MOMENTS ON A TWO-
DIMENSIONAL WING EQUIPPED WITH AN OSCIL-
LATING CIRCULAR-ARC SPOILER. Sherman A.
Clevenson and John E. Tomassoni. January 1954.
20p. diagrs., photos. (NACA RM L53K18)

Results of a wind-tunnel investigation of the forces,
moments, and phase angles on a two-dimensional
wing equipped with an oscillating circular-arc
spoiler are presented. Schlleren photographs are
presented which show the flow over and behind the
spoiler. Data for Reynolds numbers from 1.3 x 106
to 6.3 x 106, Mach numbers from 0.2 to 0 82, and
reduced frequencies from 0 to 0.92 on the normal-
force and pitching-moment coefficients and their
respective phase angles referred to spoiler position
are indicated.


11




NACA RM L53K30

PRELIMINARY INVESTIGATION OF THE FLOW IN
AN ANNULAR-DIFFUSER-TAILPIPE COMBINA-
TION WITH AN ABRUPT AREA EXPANSION AND
SUCTION, INJECTION, AND VORTEX-GENERATOR
FLOW CONTROLS. John R Henry and Stafford W.
Wilbur February 1954 27p. diagrs.
(NACA RM L53K30)

The performance of an annular-diffuser-tailpipe
combination with an abrupt area expansion was in-
vestigated with and without flow controls in the form
of suction, injection, and vortex generators The
diffuser had a 21-inch-diameter straight outer wall,
an area ratio of 1.9 to I. and fully developed pipe
flow at the inlet. Inlet Mach number was varied be-
tween 0 18 and 0.43. The ratio of the auxiliary air
flow to the flow of the main stream was varied from
0 to approximately 4 percent. IBoth suction and
injection flow controls were effective in producing
improved diffuser performance )


NACA Langley Field. V.







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