Research abstracts

MISSING IMAGE

Material Information

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
Publication Date:
Frequency:
irregular
completely irregular

Subjects

Subjects / Keywords:
Aeronautics -- Abstracts -- Periodicals   ( lcsh )
Aeronautics -- Research -- Abstracts -- Periodicals   ( lcsh )
Genre:
serial   ( sobekcm )
federal government publication   ( marcgt )
abstract or summary   ( marcgt )

Notes

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

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 001469326
notis - AGY1019
oclc - 01471285
lccn - 86657025
issn - 0499-9274
Classification:
lcc - TL501 .U5895
System ID:
AA00009235:00061

Related Items

Preceded by:
Monthly list of documents released by the NACA ...
Succeeded by:
Research abstracts and reclassification notice


This item is only available as the following downloads:


Full Text





National Advisory Committee For Aeronautics


,--Resarch Abstracts
/ ,- \


APRIL17, 1953


CURRENT IA A R POTk
D- C-

NACA TN 2792

DIRECT-READING DESIGN' -FAR
ALUMINUM-ALLOY FLAT CO l~SSI ELS
HAVING LONGITUDINAL FORMED -SECTION
STIFFENERS AND COMPARISONS WITH PANELS
HAVING Z-SECTION STIFFENERS. William A.
Hickman and Norris F. Dow. March 1953. 71p.
photos., diagrs., 8 tabs. (NACA TN 2792)

Direct-reading design charts are presented for the
determination of the stress and proportions of 24S-T3
aluminum-alloy flat compression panels having
longitudinal formed hat-section stiffeners, for given
values of intensity of loading, skin thickness, and
effective length. Hat- and Z-stiffened panels are
compared as compression panels and as the covers
of box beams.


NACA TN 2889

ESTIMATION OF HYDRODYNAMIC IMPACT LOADS
AND PRESSURE DISTRIBUTIONS ON BODIES
APPROXIMATING ELLIPTICAL CYLINDERS WITH
SPECIAL REFERENCE TO WATER LANDINGS OF
HELICOPTERS. Emanuel Schnitzer and Melvin E.
Hathaway. April 1953. 31p. diagrs (NACA
TN 2889)

An approximate method for computing water loads
and pressure distributions of lightly loaded elliptical
cylinders during oblique water impacts is presented.
The method, of special interest for the case of water
landings of helicopters, makes use of theory devel -
oped and checked for landing impacts of V-bottom
seaplanes. Comparisons of results computed by
this method with limited experimental data obtained
during drops of a circular cylinder at 00 trim and
90 flight-path angle show rough agreement. A
detailed computational procedure is included as an
appendix.


NACA TN 2905

A RAPID METHOD FOR USE IN DESIGN OF TUR-
BINES WITHIN SPECIFIED AERODYNAMIC LIMITS
Richard H. Cavicchi and Robert E. English April
1953. 72p. diagrs., 2 tabs. (NACA TN 2905)

Basic thermodynamic relations were applied to
axial-flow turbine designs for which specified aero-
dynamic limits were set in order to evolve a rapid


method of determining turbine velocity diagrams.
The method is presented in both chart and tabular
form Illustrative examples show the manner of
selecting the number of turbine stages, the optimum
work division between or among the stages, and the
proximity of the staging operation to the aerodynam-
ic limits.


NACA TN 2916

EFFECT OF THERMAL PROPERTIES ON
LAMINAR-BOUNDARY-LAYER CHARACTERISTICS.
E. B. Klunker and F Edward McLean. March1953.
29p. diagrs. (NACA TN 2916)

An iteration method is presented for solving the
laminar-boundary-layer equations for compressible
fow in the absence of a pressure gradient wherein
the temperature variation of all the fluid thermal
properties is considered. Friction and heat-
transfer characteristics have been calculated for a
stream temperature of -67o F for Mach numbers
from 1 to 10 with the use of values of the heat capac-
ity, conductivity, and viscosity determined from
experiment. Consideration of the temperature vari-
ation of all the fluid thermal properties causes the
recovery factor to decrease substantially with in-
creasing Macn numbe.-: Moreover, the heat-
transfer rate is found to be proportional to the dif-
ference between an effective enthalpy, which is a
function of both the surface temperature and stream
Mach number, and the surface enthalpy. In
contrast, the heat-transfer rate is proportional to
the difference between the recovery enthalpy and the
surface enthalpy for solutions which employ a con-
stant Prandtl number. The calculated skin friction
and heat-transfer rates based upon the use of the
Sutherland equation for viscosity and a Prandtl num-
ber of 0 75, however, are in excellent agreement
with the results of the present analysis.


NACA TN 2919

THE ASYMMETRIC ADJUSTABLE SUPERSONIC
NOZZLE FOR WIND-TUNNEL APPLICATION. H.
Julian Allen March 1953. 30p. diagrs., photos.,
2 tabs. (NACA TN 2919. Formerly RM A8E17)

An asymmetric adjustable nozzle for supersonic
wind-tunnel application which permits continuous
adjustment of the test -section Mach number is
described The characteristics of this nozzle are
compared with the more conventional supersonic
tunnel nozzles.


N4d a


*AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1794 F ST, NW. WASHINGTON 25, D C. CITING CODE NUMBER ABOVE EACH TITLE;
THE REPORT TITLE AND AUTHOR.

lo 2. d3 0


.~T






a


NACA TN 2920

INTERIM REPORT ON A FATIGUE INVESTIGATION
OF A FULL-SCALE TRANSPORT AIRCRAFT WING
STRUCTURE. M. James McGuigan, Jr. April
1953. 36p. photos., diagrs., 2 tabs. (NACA
TN 2920)

Results are presented of constant-level fatigue tests
conducted on several full-scale C-46 "Commando"
airplane wings at a level of 1 t 0. 625g or about
22 t 14 percent of the design ultimate load factor.
The average lifetime for the 34 fatigue failures that
occurred was about 200, 000 cycles. The spread
in lifetime for all failures was 4.4 to 1. 0, whereas
for failures repeatedly occurring at the same loca-
tion it was as small as 1.2 to 1.0. Effective stress
concentration factors were calculated for all failures
and indicated a value of about 4. 0 for an inspection
cutout and 2.3 for a riveted tension joint. During
the tests no change was noted in either the natural
frequency or damping characteristics of the test
specimens prior to the development of a fatigue
crack. When a crack did occur, its rate of growth
was rather slow until about 5 to 9 percent of the
tension material had failed, after which the rate of
crack growth increased rapidly.


NACA TN 2921

THE AERODYNAMIC DESIGN AND CALIBRATION
OF AN ASYMMETRIC VARIABLE MACH NUMBER
NOZZLE WITH A SLIDING BLOCK FOR THE MACH
NUMBER RANGE 1.27 TO 2.75. Paige B. Burbank
and Robert W. Byrne. April 1953. 37p. photos.,
diagrs., 5 tabs. (NACA TN 2921. Formerly
RM L50L15)

A method of designing as asymmetric, fixed geome-
try, variable Mach number nozzle has been developed
by using the method of characteristics. A small
nozzle conforming to the analytically determined or-
dinates was constructed and calibrated over a range
of Mach numbers extending from 1. 27 to 2.75. The
results show the variation in Mach number to be
0.02 or less and in the flow direction to be 0t. 20
within the test section. The range of Mach numbers
from 1. 27 to 2. 75 was obtained by translating the
lower block in a straight line parallel to the test-
section center line for a distance of 2. 17 test-section
heights.


NACA TN 2922

THE DESIGN OF VARIABLE MACH NUMBER
ASYMMETRIC SUPERSONIC NOZZLES BY TWO
PROCEDURES EMPLOYING INCLINED AND
CURVED SONIC LINES. Clarence A. Syvertson and
Raymond C. Savin. March 1953. 35p. diagrs.,
tab. (NACA TN 2922)

Two procedures are developed for designing asym-
metric supersonic nozzles for which the calculated
exit flow is essentially uniform over a range of Mach
numbers. One procedure is applicable at Mach num-


NACA
RESEARCH ABSTRACTS NO.41s


bers below approximately 3; the other procedure is
used for designs at Mach numbers exceeding 3.


NACA RM E53B17

ANALYSIS OF HEAT TRANSFER AND FLUID FRIC-
TION FOR FULLY DEVELOPED TURBULENT FLOW
OF SUPERCRITICAL WATER WITH VARIABLE
FLUID PROPERTIES IN A SMOOTH TUBE. Robert
G. Deissler and Maynard F. Taylor. April 1953.
29p. diagrs. (NACA RM E53B17)

A previous analysis of turbulent flow and heat trans-
fer for air with variable properties flowing in smooth
tubes is generalized in order to maxe it applicable to
supercritical water. The generalization is necessary
because all the pertinent properties of supercritical
water vary markedly with temperature. Calculated
velocity and temperature distributions, as well as
relations among Nusselt number, Reynolds number,
and friction factor, are presented. The effect of
variation of fluid properties across the tube on the
Nusselt number and friction factor correlations can
be eliminated by evaluating the properties at a ref-
erence temperature which is a function of both the r)
wall temperature and the ratio of wall-to-bulk
temperatures.


NACA RM L53A09

LANGLEY FULL-SCALE-TUNNEL TESTS OF THE
CUSTER CHANNEL WING AIRPLANE. Jerome
Pasamanick. April 1953. 57p. dlagrs., photos.,
tab. (NACA RM L53A09)

An investigation has been made in the Langley full-
scale tunnel to determine the lift characteristics and
some of the stability and control characteristics of an
experimental Custer Channel Wing airplane at zero
airspeed and over the low airspeed range. Compari-
son is made of the airplane static lift characteristics
as obtained in the absence of and in the presence of a
ground effect with and without the horizontal tail sur-
faces. An evaluation of the flow field about the
channel-propeller combination is included.


BRITISH REPORTS


N-11516A*

Aeronautical Research Council (Gt. Brit.)
LIST OF CURRENT PAPERS. (NOS. 51-100)
August 1952. 5p. (ARC CP 100)

This report presents a list of CP's (Nos. 51-100)
published by the Aeronautical Research Council.





NACA
RESEARCH ABSTRACTS NO.41


N-21551*

Aeronautical Research Council (Gt. Brit.)
THE NATURAL FREQUENCIES OF VIBRATION OF
PRISMATIC BLADES WITH PARTICULAR REFER-
ENCE TO A 12-STAGE TURBINE. R. Chaplin.
1952. 25p. diagrs., 3 tabs. (ARC CP 95)

The natural frequencies of vibration of the blading of
a 12 stage, 3000 rpm turbine have been measured
and compared with the values obtained by calculation.
In the calculations for the flexural modes, correc-
tions have been Introduced for shear and rotary
inertia. An empirical correction is used for the in-
fluence of the increase in torsional stiffness, due to
the platform, on the frequencies of torsional vibra-
tion. The agreement of measured and computed fre-
quencies is sufficient for the purpose of computing
critical speeds up to a frequency of five kilocycles
per second above which limit the discrepancy in-
creases with the order of the mode.


N-21552 *

Aeronautical Research Council (Gt. Brit.)
GRAPHICAL SOLUTION OF MULTHOPP'S EQUA-
TIONS FOR THE LIFT DISTRIBUTION OF WINGS.
F. Vandrey. 1952. 8p. diagrs. (ARC CP 96)

A simple graphical method is described facilitating
the determination of the lift distribution of wings.
The basis is Multhopp's method of replacing approxi-
mately the integro-differential equation for the circu-
lation by a finite system of linear equations which
give the values of the circulation at certain fixed
points along the span. The values of the unknown
circulations are represented by scales in a set of
diagrams for the equations. The multiplication of
the approximate values of the unknowns and the con-
stant coefficients of the equations is effected by
auxiliary scales in the diagrams. The corrections
of the approximate values are transferred from the
auxiliary scales to the main scales by a pair of
dividers. The lift distribution of a rectangular wing
is determined as a practical example.


N-21553 *

Aeronautical Research Council (Gt. Brit.)
FLUID DYNAMIC NOTATION IN CURRENT USE AT
N.G.T.E. S. Gray. 1952. 26p. diagrs. (ARC
CP 97)

This memorandum records and defines the current
system of notation which is in general use, at the
National Gas Turbine Establishment, for work on
axial flow compressors and cascade investigations in
general, and which is being applied to some extent to
the work on turbines. Heat transfer and supersonic
flow aspects and other specialized treatments are
excluded. Detailed definitions and explanations are
given in classified lists, illustrated by figures, and
alphabetical and numerical lists of the symbols,
suffixes and indices are included.


3


N-21554*

Aeronautical Research Council (Gt. Brit.)
SWEPT-WING LOADING. A CRITICAL COMPARI-
SON OF FOUR SUBSONIC VORTEX SHEET THEO-
RIES. H. C. Garner. FOREWORD. L. W. Bryant.
1952. 61p. diagrs., 17 tabs. (ARC CP 102)

From a systematic series of calculations of swept-
wing loading the writer has formed an opinion of the
accuracy and most useful application of vortex lattice
theory and the vortex sheet theories of Weissinger,
Multhopp and Kuchemann. The results provide a
general picture of the effect of sweep and compress-
ibility on lift slope and aerodynamic center.


N-21555*

Aeronautical Research Council (Gt. Brit.)
SOME EFFECTS OF REYNOLDS NUMBER ON A
CAMBERED WING AT HIGH SUBSONIC MACH NUM-
BERS. H. E. Gamble. 1952. 34p. photos., diagrs.,
2 tabs. (ARC CP 103)

An untapered, sweptback wing of aspect ratio 4,
sweepback 250 and section 12-percent thick (RAE
104 with 1-percent camber, a = 0.6) was tested in
the RAE high-speed tunnel. The pressure distribu-
tion was measured at the midsemispan section at
various Mach numbers up to 0.88 at Reynolds num-
bers of 0.8, 1.8 and 3. 5 x 106. There are consider-
able differences in the shapes of the pressure distri-
butions at the three Reynolds numbers, although the
boundary layer was laminar back to about 70-percent
of the chord in all cases. At the highest Reynolds
number, the suctions on the upper surface at high
Mach numbers increase from the leading-edge of the
wing right up to about 50 percent or 60 percent of the
chord, while at the lowest Reynolds number they re-
main almost constant from about 30 percent to about
60 percent of the chord. This flat-topped pressure
distribution, which is associated with a A-type shock
wave, results in lower lift coefficients and higher
pitching-moment coefficients than those obtained at
higher Reynolds number. Brief experiments, with
transition provoked near the leading edge, indicate
that the shape of the pressure distribution at
R = 0.8 x 106 is then something similar to that ob-
tained at R = 3.5 x 106 with transition free and at
about 65 percent chord. It appears however, that
inducing early transition artificially at higher
Reynolds numbers can seriously alter the pressure
distribution at high speed by causing a forward move-
ment of the shock wave and a separation aft of it.


N-21556*

Aeronautical Research Council (Gt. Brit.)
A FROST POINT HYGROMETER FOR SUPERSONIC
WIND TUNNELS. D. Beastall and A. Winyard.
1953. 8p. diagrs. (ARC CP 106)

This note describes a frost point hygrometer suitable
for measuring the water vapor content of the air in
supersonic wind tunnels at any stagnation pressure






4


within their present range of operation. It uses
CO2 as a coolant and is economical in construction
and operation.



N-21557*

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF ENDPLATES ON SWEPT WINGS.
D. Klichemann and D. J. Kettle. 1952. 23p. diagrs.
(ARC CP 104)

Existing methods of calculating the effect of end-
plates on straight wings are modified so as to apply
to swept wings. The changes in over-all lift and
drag, and also the spanwise distribution of the addi-
tional load, can be calculated. The theoretical re-
sults are compared with experimental results ob-
tained on swept wings, including new measurements
of lift, drag and pitching moment, made on an un-
tapered 450 sweptback wing of aspect ratio 3 at low
speed. The method of calculation is also extended to
cover the effect of the tip vortex which is formed on
wings without endplates.


N-21558 *

Aeronautical Research Council (Gt. Brit.)
THE DESIGN OF JETTISONABLE COCKPIT HOODS.
I. L. Keiller. 1952. 34p. photos., diagrs.
(ARC CP 105)

In the past much trouble has been experienced in the
design of jettisonable cockpit hoods and even after a
considerable amount of development many hoods are
not really satisfactory. In order that a successful
hood jettisoning mechanism can be produced it is
essential that the various problems involved should
be realized at the design stage. Consideration is
given in this paper to the jettisoning problems in-
volved in the design of all types of hoods and cockpit
covers. Certain basic design criteria are proposed
and the various methods of meeting them are dis-
cussed. Recommendations on good design practice
are given where possible. With the knowledge that is
at present available the design of a satisfactory
orthodox hood should present no great problems, but
the more advanced designs are likely to cause some
difficulty.



N-21559 *

Aeronautical Research Council (Gt. Brit.)
SOME INVESTIGATIONS INTO THE DESIGN OF
WIND TUNNELS WITH GAS TURBINE JET ENGINE
DRIVES. H. J. Higgs. 1953. 84p. diagrs. (ARC
CP 107)

The basic design of wind tunnels suitable for jet en-
gine drives has been investigated by matching the
predicted mass flows and total pressure losses of
typical tunnel configurations with the estimated mass
flow and total pressure rise characteristics of possi-
ble pumping systems. There are three possibilities


NACA
RESEARCH ABSTRACTS NO.41,


of jet engine pumping (1) induction pumping (2) suc-
tion pumping and (3) parallel suction and induction
pumping, of which induction pumping is shown to be
the most favorable. Using this system a typical
5,000 lb thrust jet engine (such as a R. R. Nene II)
can drive (1) a high speed subsonic tunnel of 4 ft2
working section up to M = 1.0, (2) a supersonic tun-
nel of 2 ft2 working section up to M = 1.2, (3) a
supersonic tunnel of 1 ft2 working section up to
M = 1.8. For these three cases it has been assumed
that humidity effects can be controlled by partial re-
circulation of the hot engine exhaust gases to raise
the working section temperature of 600 C. In the
supersonic case 600 might not be sufficient and
other methods of humidity control have been dis-
cussed in relation to jet engine drives. Induction
pumping readily allows multiple engine drive, that
is, the above working section areas can be increased
in proportion to the number of engines used.



N-21585*

Aeronautical Research Council (Gt. Brit.)
ON THE FLOW PAST A FLAT PLATE WITH UNI-
FORM SUCTION. B. Thwaites. 1952. 11p. diagrs.,
tab. (ARC R & M 2481. Formerly ARC 9391;
FM 887; Perf. 113)

A new method of performing boundary-layer calcula-
tions is introduced in this paper, and is applied to
the problem of finding the characteristics of uniform
flow past a flat plate through which there is a
constant normal velocity. An exact solution to this
problem has not yet been found and it is therefore
difficult to assess the accuracy of the results
obtained. The results, however, are compared with
those of two other methods. The new method will be
applied to other problems. When the momentum
eauation is being used, one obvious advantage of the
method is that, in "adding" velocity profiles, the
momentum thickness of each may be added to give
the momentum thickness of the whole. This is not
so in the usual methods of boundary-layer calcula-
tions, and great simplification is thereby obtained.



N-21586*

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE MAINTENANCE OF LAMINAR-FLOW
WINGS. W. E. Gray and H. Davies. 1952. 3p.
(ARC R & M 2485; ARC 10, 518. Formerly RAE
Tech. Note Aero 1862)

The maintenance of laminar-flow wings involves:
(1) the prevention of deterioration in the surface
itself ( for example, cracking of the paint or filler,
increase in roughness or waviness, etc., whether
due to weathering, stresses in flight, or accidental
damage) and (2) the prevention of contamination of
the surface with flies, etc. This report gives an
account of experience gained at the Royal Aircraft
Establishment in dealing with these problems during
flight tests on the characteristics of low-drag wings.






NACA
RESEARCH ABSTRACTS NO.41


N-21587*

Aeronautical Research Council (Gt. Brit.)
A SERIES OF LOW-DRAG AEROFOILS EMBODYING
ANEW CAMBER-LINE. OlaDouglas. 1952. 19p.
diagrs., tabs. (ARC R & M 2494; ARC 10, 620)

A series of low-drag airfoils is displayed in this pa-
per, which introduces a new stock type of camber
line. This camber line affords several advantages:
in particular, it makes it possible for any one of the
parameters CL range, PL, CLopt/Cm0 and

Merit to take values appreciably higher than can be
obtained by previous methods, the remaining param-
eters being fixed. The method of design follows
that of R & M 2166 and the greater part of the work
involves standardized computational processes. The
design of one of the series is described in detail, and
one is also compared, favorably, with a correspond-
ing roof-top airfoil. A method is given of obtaining
easily an approximate value for the lower limit of
the CL range when the upper limit is known.


N-21588*

Aeronautical Research Council (Gt. Brit.)
INVESTIGATIONS INTO THE EFFECT OF CONTIN-
UOUS SUCTION ON LAMINAR BOUNDARY-LAYER
FLOW UNDER ADVERSE PRESSURE GRADIENTS.
B. Thwaites. 1952. 23p. diagrs., tab. (ARC
R & M 2514. Formerly ARC 9555; FM 912)

The principal problem considered in this paper is
that of the flow along a surface through which fluid is
being continuously withdrawn at a constant velocity,
the velocity of the stream outside the boundary layer
having a constant and negative gradient. When no
suction is applied Howarth has solved the flow, and
the point of separation is known. With continuous
suction, the ordinary methods of using the momen-
tum equation, for example Pohlhausen's method,
breakdown, and a new method is used in this paper
to find the point of separation when suction of constant
velocity is applied. Only partial success may be
recorded in this problem, but this account of prog-
ress so far is now given in the hope that the problem
may attract other workers. The new method of
boundary-layer calculation used is not fully described
in this paper, and will be explained with examples of
its use, in a later paper. It has already been used
in R & M 2481, where a particular problem was
treated with great ease. The momentum equation
of boundary-layer flow is also used to deduce other
types of flow in which the velocity of suction or the
velocity outside the boundary layer is not constant.
These examples are inserted by virtue of the simpli-
fications of method involved, by which the momen-
tum equation may be integrated exactly; one of them
is particularly applicable to the velocity distribution
near the leading edge of a thin airfoil at high lift
coefficient, in which a small amount of suction is
sufficient to prevent the stall. There are appended
some considerations of the practical applications of
the continuous suction principle, and also some
figures showing the quantities of fluid required to be
sucked in, given for the benefit of designers who may


5



not realize how small are the quantities of fluid
sucked when separation of flow is suitably prevented.
The application of continuous suction to delay the
stall is stressed.


N-21589 *

Aeronautical Research Council (Gt. Brit.)
AEROFOIL THEORY OF A FLAT DELTA WING AT
SUPERSONIC SPEEDS. A. Robinson. 1952. 21p.
diagrs. (ARC R & M 2548; ARC 10,222. Formerly
RAE Aero 2151)

Lift, drag, and pressure distribution of a triangular
flat plate moving at a small incidence at supersonic
speeds are given for arbitrary Mach number and
aspect ratio. The values obtained for lift and drag
are compared with the corresponding values obtained
by strip theory. The possibility of further applica-
tions of the analysis leading up to the above results
is indicated.


N-21590*

Aeronautical Research Council (Gt. Brit.)
AN EXPERIMENTAL INVESTIGATION ON THE
FLUTTER CHARACTERISTICS OF A MODEL FLY-
ING WING. N. C. Lambourne. 1952. diagrs.,
photos., 6 tabs. (ARC R & M 2626. Formerly
ARC 10, 509; 0.655)

This report describes some preliminary experimental
work that has been carried out in an attempt to gain
information on the flexural-torsional flutter charac-
teristics of flying wing types of aircraft. Tests were
made with two flexible tip-to-tip models: (a) rectan-
gular plan form and (b) cranked and tapered plan
form. The method of supporting the models in the
wind tunnel allowed certain bodily freedoms to be
present either singly or simultaneously, and meas-
urements were made of critical speeds and frequen-
cies, and in a few cases the flutter motion was
analyzed by means of cinematograph records. The
experimental results are in no way conclusive and
cannot be directly applied to full-scale problems,
but they do point to some of the difficulties in the
treatment of the flutter of flying wings. Further the
difficulties encountered during the flutter tests them-
selves lead to suggested modifications in the
technique of providing a model in a wind tunnel with
the bodily freedoms appropriate to free flight condi-
tions.


N-21591*

Aeronautical Research Council (Gt. Brit.)
A CRITERION FOR THE PREVENTION OF SPRING-
TAB FLUTTER. A. R. Collar and G. D. Sharpe.
1952. 19p. diagrs., 3 tabs. (ARC R & M 2637;
ARC 9104; ARC 9956. Formerly RAE SME 3346;
SME 3378)

The present paper advances a formula which can be
used as a criterion for the degree of masp balance






6


necessary for the avoidance of spring-tab flutter.
The formula shows that if the tab is of sufficiently
light construction, mass balance may not be required
at all; on the other hand, the usual static balance
may be inadequate for a tab of high inertia. The
criterion comprehends within itself the requirement
(given elsewhere) limiting the length of a mass
balance arm. While the formula is based on theoret-
ical considerations, the numerical values for the
quantities to be used have been deduced from flight
experience, which shows excellent correlation with
the theory. Two forms for the criterion are given:
a simple form suitable for general application, and a
slightly elaborated form intended for application to
unusually large tabs. The appendix, besides contain-
ing the main analysis, also gives consideration to
certain factors which for simplicity are omitted in
the main text. In particular it is shown that the
"limiting length" for a balance arm may be general-
ized to a "limiting circle" for the position of the
balance mass: the circle can often be found from
simple geometrical considerations.


N-21592*

Aeronautical Research Council (Gt. Brit.)
PERFORMANCE CALCULATIONS FOR A DOUBLE-
COMPOUND TURBO-JET ENGINE OF 12:1 DESIGN
COMPRESSOR PRESSURE RATIO. D. H.
Mallinson and W. G. E. Lewis. 1952. 29p. diagrs.
(ARC R & M 2645; ARC 11,355. Formerly
NGTE R. 19)

This report describes a theoretical investigation
using conventional component characteristics to
discover that division of work between the low and
high-pressure compressors of a double-compound
simple-jet gas turbine of 12:1 design pressure ratio
which is likely to result in the most desirable equi-
librium operation over the normal engine speed
range. Having decided in favor of a pressure ratio
of 3:1 in the low-pressure compressor and 4:1 in the
other, a study is then made using more realistic
compressor characteristics to determine the prob-
able performance of such an engine under all flight
conditions when the design maximum temperature is
900 C (11730 K). The equilibrium running condi-
tions of the engine are investigated with special ref-
erence to the problems introduced by the double-
compound type of design.


N-21593*

Aeronautical Research Council (Gt. Brit.)
COMPRESSION TESTS ON DURAL-CELLUBOARD
SANDWICH PANELS. K. H. V. Britten. 1952.
18p. diagrs., photos., 5 tabs. (ARC R& M 2658;
ARC 10, 427. Formerly RAE Tech. Note SME 383)

Results are given of compression tests made on 56
dural-Celluboard sandwich panels with birch spruce
or whitewood centers. These are compared with
results from similar tests on dural-balsa sandwich
and all-metal panels, and it is seen that over a range
of sizes and weights considered dural-Celluboard can
be equally or more efficient for carrying end loads.


NACA
RESEARCH ABSTRACTS NO.41
-\

The birch Celluboard was more efficient than the
spruce or whitewood and the thicker sandwiches, and
those with thicker skins were more efficient than the
thinner specimens. The maximum stress reached
in the skin, 48, 000 lb/sq in., was equal to the 0.1
percent tensile proof stress of the material. The
birch filling had also reached its maximum compres-
sion stress, 8,000 lb/sq in. The design had there-
fore exploited these materials to their fullest extent.


N-21594*

Aeronautical Research Council (Gt. Brit.)
EXPERIMENTS GIVING HINGE MOMENT AND LIFT
ON A NACA 0015 AEROFOIL FITTED WITH A 40
PER CENT CONTROL, WITH ESPECIAL REFER-
ENCE TO EFFECT OF CURVATURE OF CONTROL
SURFACE. A. S. Batson, J. H. Preston and J. H.
Warsap. 1952. 27p. diagrs., photos., 6 tabs. (ARC
R & M 2698. Formerly ARC 6666; S & C 1525)

The work described in this report may be considered
a continuation of that given in part I of R & M 2008.
It may be divided into two parts. The first part con-
sists of experiments giving hinge moment and lift
data on an unmodified control forming part of an
NACA 0015 airfoil. For the second part similar
experiments were undertaken as part of a general
research into the effect on the properties of controls
of curvature of control surface.


N-21595*

Aeronautical Research Council (Gt. Brit.)
TORSIONAL VIBRATION IN AIRCRAFT POWER
PLANTS: METHODS OF CALCULATION. PART I.
INTRODUCTION AND GENERAL COMMENTS.
P,.RT II. PRACTICAL TREATMENT OF THE GEN-
ERL.L PROBLEM. PART HI. PRACTICAL CAL-
CULATIONS FOR A TYPICAL 12-CYLINDER VEE
ENGINE. B. C. Carter. 1952. 63p. diagrs.,
tabs. (ARC R & M 2739; ARC 3519. Formerly
RAE E. 3586)

The object of this report is to assist designers of
aircraft power plants in avoiding harmful torisonal
vibration of the crankshaft-airscrew system.


N-21596*

Aeronautical Research Council (Gt. Brit.)
DESIGN AND CALIBRATION TESTS OF A 5.5 IN.
SQUARE SUPERSONIC WIND TUNNEL. J.
Lukasiewicz. 1952. 36p.photos., diagrs., 4 tabs.
(ARC R & M 2745; ARC 13, 425. Formerly RAE
Tech. Note Aero 2033, sup. 100)

The main design features of the wind tunnel are
described and results are given of the investigations
carried out to determine: (a) the minimum pressure
ratio required to operate the wind tunnel at Mach
numbers up to 3. 5, and (b) the uniformity of the ve-
locity distribution in the working section at Mach
numbers of 1. 57, 1.88, 2. 48, 2.85, 3. 25 and 3. 5.
It was found that the tunnel pressure recovery can





NACA
RESEARCH ABSTRACTS NO.41


be appreciably increased by means of a contraction
("second throat") located between the working section
and subsonic diffuser. All nozzles tested were
designed with short throats and expansion profiles
with the maximum angles of expansion for the given
exit Mach number. The axial variation of Mach
number over selected intervals of working section
(not smaller than 5 in.) was found to be of the order
of +1.0 percent. It was found that condensation in
the wind-tunnel nozzle (run with atmospheric air)
has a detrimental effect on the velocity distribution
in the working section, particularly at small Mach
numbers.


N-21597 *

Aeronautical Research Council (Gt. Brit.)
INVESTIGATIONS ON STALLING BEHAVIOUR,
RUDDER OSCILLATIONS, TAKE-OFF SWING AND
FLOW ROUND NACELLES ON THE TUDOR I AIR-
CRAFT. D. J. Lyons. 1952. 18p. diagrs., 2
tabs. (ARC R & M 2789; ARC 11,412. Formerly
RAE Aero 2237)

During the development of the Tudor I aircraft, the
Royal Aircraft Establishment cooperated in the flight
tests. This report summarizes the results, which
are felt to be of general interest. The importance
of "deep tufting" in leading to an understanding of
varied aerodynamic problems has again been forcibly
demonstrated; namely in showing that: (a) early
buffeting of the Tudor as the stall is approached was
due to a very small airleak around the leading edge
of the wing root causing a breakaway of flow, the
resultant wake of which hit the tailplane; (b) early
wing-tip stalling was shown to be due to small
malfitment of the T.K. S. de-ciers; (c) rudder
"kicking" arose from flow through the hinge cutouts;
(d) excessive take-off swing was due to poor rudder
control as a result of the early rudder stall, and to
the fact that the aircraft was stalled in the ground
attitude; and (e) the inner nacelle needed consider-
able lengthening.


N-21598*

Aeronautical Research Council (Gt. Brit.)
WIND-TUNNEL MEASUREMENTS OF YAWING
MOMENT DUE TO YAWING (nr) ON A 1/5.5 SCALE
MODEL OF THE METEOR MARK F. III. J. G. Ross
and R. C. Lock. 1952. 31p. diagrs., photos., 10
tabs. (ARC R & M 2791; ARC 10, 786. Formerly
RAE Aero 2199)

During recent investigations into the self-excited
oscillations in yaw, experienced on Meteor aircraft,
the lateral stability derivative nr was measured
in flight, and found to differ considerably during
initial experiments from the theoretical estimate. A
new technique was therefore devised to measure nr
in the wind-tunnel; and, with its aid, modifications
Were tested on a model with the object of reducing
the self-excited oscillations in flight. Measure-
ments of nr were made over a range of Reynolds
numbers, and for different periods of oscillation of
the model. The final comparison of the flight and


7


wind-tunnel tests, after certain refinements in
technique of the former, and after corrections for
solid friction to the latter had been made, showed
that the full-scale measurement of nr was about
10 percent less than that obtained in the tunnel.
Considering the difficulties involved, this agreement
may be considered as satisfactory. For the model
in the standard condition, the value of nr was about
20 percent less than the estimated figure of -0. 108
at zero lift, but with dorsal fins. It was found
possible, without altering the value of nv, to
increase the value of nr to the estimated value.
The "snaking" tendencies of the model, which were
more pronounced at small angles of incidence,
could be greatly reduced by fitting an upper dorsal
fin.


MISCELLANEOUS


NACA TN 2505

Addendum No. 1 to "ON THE ATTACHED CURVED
SHOCK IN FRONT OF A SHARP-NOSED AXIALLY
SYMMETRICAL BODY PLACED IN A UNIFORM
STREAM." S. F. Shen and C. C. Lin. October
1951.


NACA-Langley 4-17-53 4000




UNIVERSITY OF FLORIDA
31 262081l ll111111 5 4111111 -
3 1262 08153 285 4




Full Text
xml version 1.0 encoding UTF-8
REPORT xmlns http:www.fcla.edudlsmddaitss xmlns:xsi http:www.w3.org2001XMLSchema-instance xsi:schemaLocation http:www.fcla.edudlsmddaitssdaitssReport.xsd
INGEST IEID E4OTLTFLF_X7UJ9H INGEST_TIME 2012-02-29T18:20:01Z PACKAGE AA00009235_00061
AGREEMENT_INFO ACCOUNT UF PROJECT UFDC
FILES