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
Publication Date:
Frequency:
irregular
completely irregular

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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:
AA00009235:00004

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Full Text

20000" F IN 2- AND 4I-INCH-DIAMETER HORIZON-
TAL TUBES. George R. Kinney, Andrew E
Abramson and John L. Sloop. 1952. ii. 21p.
diagrs photos 5 tabs. INACA Rept 10871.
Formerly RM E50Fl9: RM E51Cl3; RM E52B20)

An experimental investigation of internal-lIquid-film
cooling was conducted In 2- and 4-lnch-dramleter
straight metal tubes with air flows at 60)00 to 20000 F
and Reynolds numbers from 2. 2 to 14 x 105 The
coolant was water at flows of 0. 8 to 12 percent of
air flow. Visual observations of liquid-film flows
were made In transparent tubes with air flows at 800
and 8000 F and diameter Reynolds numbers from 4. I
ro 29 x 105. Flows of water, water-detergent solu-
tlons, and aqueous ethylene glycol solutions were
rnvestigated. Liquad-coolant films were established
and maintained around and along the tube wall In
..---concurrent flow' with the hot alr. The tube w~all wras
40 h'EPR below the boiling temperature of the coolant
RNhe surfaces covered by liquid coolant. Coolant
fi re relatively smooth unless the coolant flow
waS\ iclently; high so that the liquid film was
i gy dabugh to enter the region where turbulent
forces predominate over viscous forces; wavelike
disturbances then developed on the liquid film.
T se disturbances resulted In Increased loss of
tcoolant from the film and reduced effectiveness of
the coolant.

NACA Rept. 1093

HEAT TRANSFER TO BODIES IN A HIGH-SPEED
RAREFIED-GAS STREAM. Jackson R. Stalder,
Glen Goodwin and Marcus O. Creager 1952. 1I,
10p diagrs., tab. (NACA Rept. 1093 Formerly
TN 2438)

The equilibrium temperature and heat-transfer
coefficients for transverse cylinders in a high-speed
stream of rarefied gas were measured over a range
of Knudsen numbers from 0. 025 to 11.8 for Mach
numbers from 2.0 to 3.3. Fully developed free-
molecule flow was found to first occur at Knudsen
numbers of approximately 2. 0. The Nusselt number
was found to be a function of the Reynolds number
only, and the temperature-reov~ery factor to depend
primarily on the Knudsen number.

NACA TN 2908

DETERMINATION OF MEAN CAMBER SURFACES
FOR WINGS HAVING UNIFORM CHORDWISE
LOADING AND ARBITRARY SPANWISE LOADIN'G
TN SUBSONIC FLOW. S. Katzoff, M. Frances
Falson and Hugh C. DuBose. May 1953. 43p
diagrs., tab (NAC'A TN 2908)


CURRENT NACA REPORTS

NACA Rept. 1074

HYDRODYNAMIC IMPACT OF A SYSTEM WITH A
SINGLE ELASTIC MODE. I THEORY AND
GENERALIZED SOLUTION WITH AN APPLICATION
TO AN ELASTIC AIRFRAME. Wilbur 1,. Mayo.
1952. ii, 17p. diagrs., 2 tabs. (NACA Rept. 1074.
Formerly TN 1398)

Solutions of impact of a rigid prismatic float con-
nected by a massless spring to a rigad upper mass
are presented. The solutions are based on hydrtody-
namic theory which has been experimentally con-
firmed for a rigid structure. Equations are given for
defining the spring constant and the ratso of the
sprung mass to the lower mass so that the two-mass
system provides representation of the fundamental
mode of an airplane wing. The forces calculate ae
more accurate than the forces which would be ,'
dieted for a rigid airframe since the effect of
fundamental mode on the hydrodynamic force t
into account. The response of the two-mass t
gives the response of the represented mode ;
although no provision is made for taking into account
the effect of secondary modes on the hydrodynamic
force, means are indicated whereby the results may
be used to approximate the response of modes other
than the fundamental mode.


NACA Rept. 1076

EFFECTS ON LONGITUDINAL STABLLITY AND
CONTROL CHARACTERISTICS OF A BOEING B-2g
AIRPLANE OF VARIATIONS IN STICK-FORCE AND
CONTROL-RATE CHARACTERISTICS OBTAINED
THROUGH USE OF A BOOSTER IN THE ELEVATOR-
CONTROL SYSTEM. Charles W. Mathews, Donald
B. Talmage and James B. Whatten. 1952. ii, 17p.
diagrs., photo., tab. (NACA Rept. 1076. Former-
ly TN 2238; RM L50D11)

The longitudinal stability and control characteristics
of a B-29 airplane have been measured with a con-
trol surface booster Incorporated an the elevator-
control system. The measurements were obtained
with the booster operating to provide various control-
force gradients and various maximum rates of con-
trol motion. Results are presented whrch show the
effect of these booster parameters on the handling
qualities of the test airplane.

NACA Rept. 1087

INTERNAL -LIQUID- FLM-COOLING EXPERI-
MENTS WITH AIRSTREAM TEMPERATURES TO
'AVAILABLE ON LOAN ONLY


ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1744 F ST., NW, WASHINGTON 1s. D C., CITING CODE NUMBER ABOVE EACH TITLE:
THE REPORT TITLE AND AUTHOR


National Advisory Committee For Aeronautics


Research Abstracts


NO. 44


JUNE 12, 1953






NACA
RESEARCH ABSTRACTS NO.44

of: the enamel into pits formed by galvants att ck of
the enamel on the steel, was first examined frm
the standpoint of data on adherence obtained In
earlier studies. Also, several exp~eriments were
performed which demonstrated that galvanic corro-
sion could occur during the short firing times
encountered in enamel processing. However,
inconsistencies were observed in the data which
indicated that the mechanism, of galvanic attack
followed by mechanical anchoring was not the only
important factor affecting bond st rength.

NACA TN 2941

THE DRAG OF FINIE-LENGTH~ CYLINDERS
DETERMINED FROM FLIGHT TESTS AT HIGH
REYNOLDS NUMBERS FOR A MACH NUMBER
RANGE FROM 0. 5 TO 1. 3. Clement J. Welsh.
June 1953. 12p.diagrs., photo. (NACA TN 2941.
Formerly RM L52C31)

Results of a free-flight investigation to determine
the drag of circular, finite-length cylinders are
presented for a Mach number range from about 0. 5
to 1.3. Also included are drag results of previous
extperimental tests of Infinite-length cylinders. Drag
of circular cylinders at supersonic speeds is largely
independent of fineness ratio and Reynolds number;
whereas, at subsonic speeds, the drag of famtle-
length cylinders (fineness ratios of about 60 and
below) increases as their fineness ratios increase.

NACA TN 2942

PRESSURE DISTRIBUTIONS ABOUT FINTE
WEDGES IN BOUNDED AND UNOUND]ED SUBSONIC
STREAMS. Patrick L. Donoughe andErnst I.
Prasse. May 1953. 41p. diagrs., photos., 2 tabs.
(NACA TN 2942)

An analytical investigation of incompressible flowr
about wedges was made to determine effects of
tunnel-wedge ratio and wedge angle on the wedge
pressure distributions. The region of applrcability
of infinite wedge-type velocity distribution was
examined for finite wedges. Theoretical and exp~eri-
mental pressure coefficients for various tunnel-
wedge ratios, wedge angles, and subsonic Mach numt-
bers were compared.


NACA TN 2043

THE ATTENUATION CHARACTERISTICS O1F FOUR
SPECIALLY DESIGNED MUFFLERS TESTED ON A
PRACTICAL ENGINE SETUP. George Mi. Stokes
and Don D. Davis, Jr May 1953. 30p. diagrs.,
photos., tab. (NACA TN 2943)

Attenuation characteristics of four different
resonator mufflers wrere determined in both cold
tests and engine field tests and compared with the
theoretical calculations. These mufflers were
specifically designed for a helicopter. Engine-
exhaust sound pressures, temperatures, and noise
levels from the helicopter were measured. The
experimental muffler cold tests indicated close a
agreement with theory, whereas the engine tests


Methods involving integration around the wing
boundary are presented for computing mean camber
surfaces to support uniform chordwise loading, with
either uniform or nonuniform spanwise loading, in
subsonic flow. For polygonal wings with uniform
area loading, analytical expressions are developed
for both the slopes and the ordinates of the mean
chamber surfaces. Mean camber surfaces of several
calculated wings are shown.

NACA TN 2924

COMBINED-STRESS FATIGUE STRENGTH OF
76S-T61 ALUMINUM ALLOY WITH SUPERIMPOSED
MEAN STRESSES AND CORRECTIONS FOR
YIELDING. William N. Findley, University of
Illinois. May 1953. 90p. diagrs., photos., 3 tabs.
(NACA TN 2924)

Fatigue data for 76S-T61 aluminum alloy are
presented for several combinations of bending and
torsion with both alternating and mean stresses.
Special fatigue equipment for these tests is
described. Correction for yielding was applied to
the spean stresses in the bending and in the torsion
fatigue tests. The literature on the effect of com-
bined stress and of mean stress in fatigue is '
reviewed and the results of the present series are
compared with those of various theories. A new
notation for state of stress and a new criterion for
combined stress fatigue are discussed. A correction
for anisotropy is proposed and its effect discussed.
Energy theories of fatigue under combined stress are
critically examined and a test proposed. The
observed mode of fracturing is described and a
qualitative theory of the mechanism of formation and
propagation of fatigue cracks is proposed.

NACA TN 2930

STRENGTH ANALYSIS OF STIFFENED THICK
BEAM WEBS WITH RATIOS OF WEB DEPTH TO
WEB THICKNESS OF APPROXIMATELY 60*
L. Ross Levin. May 1953. 11p. photo., diagrs. >
2 tabs. (NACA TN 2930)

The results of an experimental investigation of the
strength of plane diagonal-tension webs with ratios
of web depth to web thickness of about 60 are pre-
sented. An analysis of the beams indicated that the
methods of strength analysis presented in NACA TN
2661 are applicable to beams with the flanges
symmetrically arranged with respect to the web if
the portal-frame effect is taken into account.

NACA TN 2935

THE GALVANI CORROSION THEORY FOR
ADHERENCE OF PORCELAIN-ENAMEL GROUND
COATS TO STEEL. D. G. Mloore, J. WT. Pitts,
J. C. Richmond and W. N. Harrison, National
Bureau of Standards. June 1953. 19p. diagrs.,
photos. (NACA TN 2935)

The galvanic corrosion theory of adherence between
ground-coat enamels and steel was investigated-
The theory, which is based on mechanical anchoring






NACA
RESEARCH ABSTRACTS NO.44


indicated some discrepancies. Test results show
the usefulness of the theoretical equation used for
predicting muffler attenuation characteristics.


NACA TN 2945

THE CREEP OF SINGLE CRYSTALS OF
ALUMINUM. R. D. Johnson, F. R. Shober and
A. D. Schwope.~Battelle Mlem~orial institute. May
1953. 51p. diagrs., photos., tab. (NACA TN 2945)

The creep of single crystals of high-purity aluminum
was investigated in the range of temperatures from
room temperature to 4000 F and at resolved-shear-
stress levels of 200, 300, and 400 psi. The tests
were designed in an attempt to produce data regard-
ing the relation between the rate of strain and the
mechanism of deformation. The creep data are
analyzed in terms of shear strain rate and the results
are discussed with regard to existing creep theories.
Stress-strain curves were determined for the
crystals in tensil and constant-load-rate tests in the
same tempe rature range to supplement the study of
plastic deformation by creep with information
regarding the part played by crystal orientation,
differences in strain markings, and other variables
in plastic deformation.

NACA TN 2946

A SMALL PIRANI GAGE FOR MEASUREMENTS OF
NONSTEADY LOW PRESSURES. M. John Pilny.
June 1953. 1, 36p. diagrs., photos. (NACA
TN 2946)

A small Pirani gage made of surgical grain-of-
wheat lamps and its operating equipment for the
measurement of 12 channels of low pressures in the
range 0. 1 to 10 mm Hg abs are described, together
with techniques of calibration and use. Theoretical
and measured characteristics of small Pirani gages
are presented to show the suitability.,of such gages
for the precise measurement of nonsteady low
pressures.


NACA TN 2947

A VISUALIZATION STUDY OF SECONDARY FL;OWS
IN CASCADES. Arthur G;. Hansen, Howard z*
Herzig nd George R. Costello. May 1953. 93p-
photos., diagrs. (NACA TN 2047. Formerly
RM E52FL9)

Flow-visualization techniques are employed to
ascertain the streamline patterns of the
nonpotential, secondary flows in the boundary layers
of cascades, thereby providing a basis for more
extended analyses in turbomachines. The three-
dimensional deflection of the end-wall boundary
layer results in the formation of a vortex well up in
each cascade passage. The size and tightness of the
vortex generated depend upon the main flow turning
in the cascade passage. Once formed, a vortex
resists turning in subsequent blade rows. This
results in unfavorable angles of attack and possible
flow disturbances on the pressure surfaces of sub-


sequent blade rows when the vortices impinge on
these surfaces. Two major tip-clearance effects
are observed: the formation of a tip-clearance
vortex, and the scraping effect of a blade! with rela-
tive motion past the wall boundary layer. The flow
patterns indicate methods for improving the blade-
tip loading characteristics of compressors and of
low- and high-speed turbines.


NACA TN 2949

A VARIABLE-FREQUENCY LIGHT SYNCHRONIZED
WITH A HIGH-SPEED MOTION-PICTURE CAMERA
TO PROVIDE VERY SHORT EXPOSURE TIMES.
Walter F. Lindsey and JosephBurlock. May 1953.
17p. photos., diagrs. (NACA TN 2940)

A new high-speed photographic technique has been
developed which employs a variable-frequency light
synchronized with a commercially available 16-
millimeter high-speed mnotion-picture camera with-
out appreciable alterations to the camera. The
technique is described and results obtained by this
technique of photographing the flow past models in a
wind tunnel employing the schlieren method of flow
visualization are presented. The photographs show
that the new technique, through the use of extremely
short exposure times (about 4 microseconds),
provides more sharply defined pictures throughout
the flow field than were obtained by conventional.
techniques.


NACA TN 2950

A NEW SHADOWGRAPH TECHNIQUE FOR OBSER-
VATION OF CONICAL FLOW PHENOMENA IN
SUPERSONIC FLOW AND) PRELIMINARY RESULTS
OBTAINED ~FOR TRIANGULAR WING. Eugene
S. Love and Carl E. Grigsby. May 1953. 16p.
diagrs., photos. (NACA TN 2950)

A new shadowgraph technique is presented for the
observation of conical flow phenomena in supersonic
flow in a plane normal or nearly normal to the axis
of propagation. Preliminary results obtained for a
triangular wing are also presented.

NACA TN 2951

FLIGHT INVESTIGATION OF THE EFFECT OF
TRANSIENT WING RESPONSE ON WING STRAINS
OF A FOUR-ENGINE BOMIBIER AIRPLANE IN
ROUGH AIR. Harold N. Murrow and Chester B.
Payne. June 1953. 24p. diagrs., 2 tabs.
(NACA TN 2951)

The results of a flight investigation on a four-engine
bomber airplane to determine the effect of wing
felebiity on wing strains show that the average
amplification at the wing root station for bending
strains per g in gusts was approximately 31 percent
higher than the bending strains per g in slow pull-
ups. The amplification factor decreased in the
outboard direction, except for the most outboard
station where it increased slightly. At the inboard
stations the vibratory inertia effects on spar shear
strains appear to induce the primary strains. At








4

the outboard stations the amplification factors for the
shear strains were approximately the same as those
for the bending strains.

NACA TN 2952

IMPINGEMENT OF WATER DROPLETS ON NACA
651-208 AND 651-212 AIRFOILS AT 40 ANGLE OF
ATTACK. Rinaldo J. Brun, Helen M. Gallagher
and Dorothea E. Vogt. May 1953. 49p. diagrs.
(NACA TN 2952) '

The trajectories of droplets in the air flowing past
an NACA 651-208 airfoil and an NACA 651-212
airfoil, both at an angle of attack of 40, were
computed with a mechanical analog. The amount
of water in droplet form impinging on the airfoils,
the area of droplet impingement, and the rate of
droplet impingement per unit area on the airfoil
surface affected were calculated from the trajec-
tories.


NACA TN 2953

AN INVESTIGATION OF THE EXPERIMENTAL
AERODYNAMIC LOADING ON A MODEL HELICOP-
TER ROTOR BLADE. John R. Meyer, Jr. and
Gaetano Falabella, Jr., Massachusetts Institute of
Technology. May 1953. 110p. diagrs., photos.
(NACA TN 2953)

Pressure distributions were measured on a model
helicopter rotor blade under hovering and simulated
forward-flight conditions. Pressures were recorded
at advance ratios of 0. 10, 0. 22, 0. 30, 0. 40, and
0. 50 for a zero-offset flapping-hinge rotor and at
0. 10, 0.22, 0.30, 0. 45, 0.60, 0.80, and 1.0 for a
lifting rotor having a flapping-hinge offset of 13
percent. Analysis of angle of attack at the tip of the
retreating blade indicated that an appreciable offset
flapping hinge in combination with a low blade mass
constant offers a means of postponing stall on the
retreating blade.

NACA RM 53D14

EFFECTS OF HIGH DEGREES OF BIAXIAL
STRETCH-FORMING ON CRAZING AND OTHER
PROPERTIES OF ACRYLIC PLASTIC GLAZING.
I. Wolock, B. M. Axilrod and M. A. Sherman.
National Bureau of Standards. May 1953. 19p.
photos., diagrs., 3 tabs. (NACA RM 53D14)

Tests were conducted on sheets of commercial cast
polymethyl methacrylate that were hot-stretched
approximately 100 and 150 percent and on
unstretched control material. Tensile strength
increased only for the 150-percent-stretched
material. Total elongation was higher for the
stretched materials than for the unstretched but
decreased as the degree of stretching increased.
None of the stretched specimens crazed in the short-
time tensile tests. The threshold stress for stress-
solvent crazing with benzene increased from approxi-
mately one-fourth of the ultimate strength for the
unstretched material to approximately three-fourths
for the 100-percent-stretched material. Most of
the 150-percent-stretched specimens did not solvent


NACA
RESEARCH ABSTRACTS N0.44

craze at stresses very close to the ultimate strength.
Abrasion resistance of the stretched material'as
appreciably less than that of the unstretched
material.

NACA RM E53C26

DE-ICING AND RUNBACK CHARACTERISTICS OF
THREE CYCLIC, ELECTRIC, EXTERNAL DE-ICING
BOOTS EMPLOYING CHORDWISE SHEDDING.
Robert S. Ruggeri. May 1953. 32p. photos.,
diagrs. (NACA RM E53C26)

The performance characteristics of three cyclic,
electric, rubber-clad de-icing boots were evaluated.
Each boot was operated in icing at design specifica-
tions of 21 watts per square inch for cycled areas,
13 watts per square inch for continuously heated
parting strips, a heat-on time of 10 seconds, and a
cycle ratio of 10. For a free-stream velocity of
approximately 395 feet per second, the range of free-
stream total temperature at which the icing protec-
tion afforded by the various boots became marginal
was from 120 to 150 F for values of liquid-water
content employed. The runback characteristics
of the boots were similar. The forward cycled
segments, upper and lower surfaces, were the most
critical areas for the three boots investigated.


NACA RM L53A30

GUST-TUNNEL INVESTIGATION TO DETERMINE
EFFECTS OF CENTER-OF-GRAVITY POSITION ON
THE GUST LOADS OF A DELTA-WING MODEL
WITH LEADING EDGE SWEPT BACK 600. Thomas
D. Reisert and Domenic J. Maglieri. June 1953.
13p. diagrs., photo., 2 tabs. (NACA RM L53A30)

A gust-tunnel investigation to determine the effect
of center-of-gravity position on the gust loads on a
delta-wing model with the leading edge swept back
600 was made over a range of center-of-gravity
positions from 4 percent ahead of to 11 percent
behind the leading edge of the mean geometric chord.
This investigation indicated that a 1-percent rear-
ward movement of the center-of-gravity position
increased the acceleration increment in the sharp-
edge gust by 0. 5 percent. In a gust with a gradient
distance of 6. 5 chords, the acceleration increment
was increased by approximately 2 percent for the
same movement of the center of gravity. Compari-
son of these results with those for a conventional
airplane model indicates that the change in load for
both configurations would be nearly the same in a
sharp-edge gust, but in a gust with a gradient dis-
tance of 6. 5 chords, the change in load for a given
change in center-of-gravity position would be
approximately twice as great for the delta-wing
model as for the conventional airplane model.


NACA TM 1350

THE MICROSTRUCTURE OF TURBULENT FLOW.
(Mikrostructura turbulentnogo potoka.) A. M.
Obukhoff and A. M. Yaglom. June 1953. 41p.
diagrs. (NACA TM 1350. Trans. from Prikladnaya
Matematika i Mekhanika, v. 15, 1951, p.3-26).






NACA
RESEARCH ABSTRACTS NO.44

An attempt is made to describe quantitatively the
structure of the velocity, pressure, and accelera-
tion fields for all scales for which the theory of
Kolmogoroff is applicable.


BRITISH REPORTS


N-21801*

Aeronautical Research Council (Gt. Brit.)
THE METHOD OF INFLUENCE FACTORS IN
ARITHMETICAL SOLUTIONS OF CERTAIN FIELD
PROBLEMS. PART I. A. Thom. PART II.
Laura Klanfer. 1953. 30p. diagrs., tab. (ARC
R & M 2440. Formerly ARC 9854; FM 968;
ARC 11,010; FM 968a)

This paper gives an extension to the squares method
of solving certain field problems. The idea of
influence factors is used. This enables certain
problems to be solved without squaring and reduces
the actual arithmetical work in others to a fraction
of that otherwise required. As an example, the
effect of the channel walls is calculated for the flow
of a compressible fluid past given profiles in chan-
nels of different widths. A comparison is also
made of the resulting blockage factor with the values
given by the mass-flow method described in previous
work.


N-21802*

Aeronautical Research Council (Gt. Brit.)
TRANSITION AND DRAG MEASUREMENTS ON THE
BOULTON PAUL SAMPLE OF LAMINAR-FLOW
WING CONSTRUCTION. Part 1: MEASUREMENTS
IN THE 13 x 9 FT TUNNEL AT THE N. P. L. J. H.
Preston and N. Gregory. Part II: MEASURE-
MENTS IN THE NO. 2, ll-FT TUNNEL AT THE
R. A. E. K. W. Kimber. Part III: DISCUSSION OF
RESULTS, AND NOTE. J. H. Preston. 1952.
23p. photos., diagrs., tab. (ARC R & M 2499;
ARC 10, 123. Formerly RAE Aero 2170)

Transition tests and drag measurements have been
carried out on the Boulton Paul sample of laminar-
flow wing construction up to R = 9 x 106 in the
National Physical Laboratory 13- by 9-foot tunnel
and in the Royal Aircraft Establishment 11-1/2 by
8-1/2 foot tunnel up to R = 15 x 106. At small
incidences lying between 20 and at Reynolds num-
bers up to 9 x 106, the agreement between measure-
ments made in both tunnels is good. At incidences
of t30 and t40 the transition moves forward with
increase of speed more rapidly in the R. A. E.
tunnel than in the N. P. L. tunnel, and the transition
front is considerably more irregular in the R. A. E.
tunnel. This difference occurs in spite of the
appreciably less measured turbulence of the R. A. E.
tunnel, which may be expected to show up, on a wing
with appreciable waviness, near the limit of the low-
drag range. In the N. P. L. tunnel the theoretical
low-drag CL range of tD. 35 is maintained up to
about R = 7 x 106 with transition back to beyond
0. 4c (pressure minimum at 0. 45c). In the R. A. E.
tunnel, the corresponding R is slightly less. The


5

low-drag range decreases as R is increased until,
at R = 15 x 106 in the R. A. E. tunnel, the low-drag
range of incidence is only 10. This reduction is
attributed to waviness of the model, exaggerated by
the turbulence of the tunnel stream. Nevertheless
the wing, considering its waviness, has performed
remarkably well in both tunnels, especially when
compared with previous models of lower waviness.


N-23040*

Aeronautical Research Council (Gt. Brit.)
SOME EXPERIMENTS ON THE HEAT TRANSFER
FROM A GAS FLOWING THROUGH A CONVERGENT-
DIVERGENT NOZZLE. 0. A. Saunders and P. H.
Calder. January 5, 1952. 13p. diagrs., 2 tabs.
(ARC 14, 539; R. 222)

Heat transfer at high subsonic and supersonic speeds
is more complicated than at normal velocities
because the cross-sectional temperature distribution
is not usually fully developed. The central core of
gas is often unaffected by the heat added, in contrast
to one-dimensional flow calculations which assume
the heat to be spread over the whole cross section.
Experiments are described using a water-cooled
convergent divergent nozzle of smooth continuous
profile through which hot gases at 8650 C were passed
and the heat transfer measured at different positions
along the divergent portion, at Mach numbers up to
1. 75. The results are very consistent when clotted
in terms of the length Reynolds number measured
from the throat. Velocity traverses at the exit also
confirm that the boundary layer may be assumed to
be turbulent and to commence at the throat. The
heat-transfer results also agree very well with the
formulas for turbulent flow over a flat plate at low
speeds, suggesting that such low spee4 formulas
may be used for supersonic flow in nozzles. Some
tests with a straight pipe at high subsonic speeds
give results somewhat higher than the flat plate
formula, due probably to pipe radius effects.

N-23041*

Aeronautical Research Council (Gt. Brit.)
THE TURBULENT BOUNDARY LAYER IN
DIFFUSING DUCTS. (Abbreviated thesis) R. R.
Shaw. January 17, 1952. 42p. diagrs., tabs.
(ARC 14, 587; FM 1667; Oxford Univ., Engineering
Lab. No. 57)

The work reported in this thesis has been primarily
a study of the flow in the (turbulent) boundary layer,
but with particular emphasis on those aspects of
this subject which are pertinent to calculations in
a conical diffuser, and the thesis concludes with the
application of the results obtained to a few typical
calculations on a diffuser. The work has been
restricted to the cone of an incompressible fluid,
and to the turbulent (as distinct from the laminar)
boundary layer.








NACA
RESEARCH ABSTRACTS NO.44


past cones. Semicone angles of 150, 200, 22. 50, .
250, and 300 are considered, and the range of free
stream Mach number is approximately 1. 3 to 8. 0.


N-23154*

Aeronautical Research Council (Gt. Brit. )
THE PRESSURE ON THE: SURFACE OF A FLAT
ELLIPTIC CONE SET SYMMETRICALLY IN A
SUPERSONIC STREAM. D. G. Hurley. 1953.
18p. diagrs., tab. (ARC CP 109)

The first order solution of the problem, of the
supersonic flow past a flat elliptic cone set
symmetrically to the wind indicates that the
pressure over the surface is constant if the body
lies within the Mach cone of the apex. This result
is incorrect near the leading edges of the cone and
an improved solution is derived here by the introduc-
tion of line sources near the leading edges.
Numerical results are given for three bodies.

N-23155*

Aeronautical Research Council (Gt. Brit.)
A NOTE ON SHOCK TUBES. D. W. Holder. 1953.
13p. diagrs. (ARC CP 110)

A brief description of the flow in shock tubes is
given and some of the problems which have been
investigated by using shock tubes are considered.

N-23156*

Aeronautical Research Council (Gt. Brit.)
TAILPLANE LOADS AND NORMAL ACCELERA-
TIONS AFTER AN AUTOMATIC CONTROL FAIURE.
J.L. Reddaway. 1953. 36p. diagr., 2tabs,
(ARC CP 111)

A method is given for the calculation of aircraf
behavior when failure of the autopilot in pitch
produces sudden elevator movements. Expressions
for the changes in aircraft normal acceleration, tail
unit accelerations and aerodynamic tail loads are
derived. These expressions together waith their
maxima and the times taken to reach these m~axm
are tabulated. A calculation on specifics aircraft
shows the use of the table.


N-23157*

Aeronautical Research Council (Gt. Brit. )
LOW SPEED WIND TUNNEL INVESTIGATION OF
TAB HINGE MOMENT CHARACTERISTICS.
W. J. G. Trebble and J. F. Holford. 1953. 20p.
diagrs., 4 tabs. (ARC CP 112)

Hinge moments have been measured on tabs of 4. 7-
percent local chord on a tailplane with 14o trailing-
edge angle. The range of investigation covered the
effects of 32-percent nose balance and of the gap
between tab and elevator. For small deflections of
the control surfaces cl and c2 are negligible while
e3 is -0. 36 and -0. 28 for the unaanced and
balanced tabs, respectively. With large angles of


N-23043 *

Aeronautical Research Counnil (Gt. Brit.)
SOME DEFICIENCIES IN CURRENT
METEOROLOGICAL KNOWLEDGE IN RELATION
TO THE OPERATION OF JETr AIRCRAFT. P. A.
Shp Tr. ,eray0 A1 4). 3p. (ARC 14,709,

Meteorological problems which have emerged in the
design and operation of jet aircraft are discussed
and the need for further knowledge in this field is
pointed out.


N-23121 *

National Gas Turbine Establishmnent(Gt. Brit. )
THE BURNING OF SINGLE DROPS OF FUEL. Part
V. THE COMBUSTION RATES OF SOLID FUEL
PARTICLES. G. A. E.Godeave. November 1952.
35p. diagrs., 5 tabs. (NGTE R. 126)

A method of calculating the burning rates of solid
spherical particles is proposed, using as a basis
Friissling's semiempirical equation for the transfer
of mass in the neighborhood of spherical surfaces.
The theoretical treatment shows good agreement
with published experimental results on the burning
rates of carbon spheres. This being so, the
theoretical equation may be used to calculate in
particular the effects of Reynolds number and partial
pressure of oxygen on the combustion rates.

N-23122 *

National Gas Turbine Establishment (Gt. Brit. )
OUTLINED GENERAL TREATMENT OF THE
CALCULATION OF WAVE EFFECTS DUE TO
SMALL, DISTURBANCES OF STEADY STABLISED
BURNING. PART I. P. W.H. Howe. November
1952. 51p. diagrs. (NGTE R. 128)

A procedure is outlined for calculating the transient
effects caused by typical disturbances of steady
burning conditions, in a duct employing a stabilized
flame and a fuel with any burning characteristics.
The method of solution is to express the available
information as analytical formulas applying to
simple cases, and then to build up an iterative
process using the simple cases.

N-23153*

Aeronautical Research Council (Gt. Brit.)
SOME CURVES FOR USE IN CALCULATIONS OF
THE: PERFORMANCE OF CONICAL CENTREBODY
INTAKES AT SUPERSONIC SPEEDS AND AT FULL
MASS FLOW. L. E. Fraenkel. 1953. 33p.
diagrs., 2 tabs. (ARC CP 108)

This note presents curves of a number of parameters
which frequently occur in calculations of the
performance of conical centerbody intakes at super-
sonic speeds and full mass flow. These parameters
follow directly from the theoretical supersonic flow






NACA
RESEARCH ABSTRACTS NO.44


the elevator, and with moderate angles when the
elevator gap is open, eg tends to the calculated
value for this wing without boundary-layer terms.
Consequently the enrve of tab hinge moment as the
tab and elevator both move is not linear. Values of
cl and c3 calculated by thick airfoil theory, with
Bryant's empirical boundary-layer terms, are mn
good agreement with measured values for small
deflections.

N-23158*
Aeronautlcal Research Council (Gt. Brit.)
AN APPROXIMATE METHOD OF DERIVING THE
TRANSIENT RESPONSE OF A LINEAR SYSTEM
FROM THE FREQUENCY RESPONSE. C. A. A.
Wass and E.G. Hayman. 1953. 20p. diagrs. >
2 tabs. (ARC CP113)

There are several published methods of deriving the
transient response of a linear system from the known
frequency response, but these are all rather lengthy'
Another method is described, which is much more
rapid, although less accurate. It is based on the
calculation of the response of the system to a square
wave as expressed by a Fourier series. For any
system, there is an optimum square wave frequency,
and the process of selection of this fundamental
frequency is described. It is shown that considera-
tion of responses up to the eleventh harmonic only
can give transient response curves which are in
error by less than 2 percent. A description is given
of a circular computer which speeds the calculations,
and twoe tables of values are included for use with
the computer.

31-23159.

Aeronautical Research Council (Gt. Brit. )
A METHOD FOR CALCULATING PRESSURE
DISTRIBUTIONS ON CIRCULAR ARC OGIVES AT
ZERO INCIDENCE AT SUPERSONIC SPEEDS,
USIG THE PRANDTIL-MEYER FLOW RELATIONS.
H. K. Zienkiewicz. 1953. 12p. diagrs., tab.
(ARC CP 114)

A method is given for a rapid determination of
pressure distributions on circular are ogives at zero
incidence at supersonic spe~eds. The method is
based on the characteristics pressure distributions
and involves the use of the two-dimnensional Prandtl-
Meyer flow relations. It gives very good agreement
with characteristics results from other sources and
with experimental results. The method can
probably be extended to ogives with profiles other
than the circular are.


N-23160*

Aeronautical Research Council (Gt. Brit.)
CHARS OF THE WAVE DRAG OF WINGS AT ZERO
LIFT. T. Lawrence. 1953. 22p. diagrs.
(ARC CP 116)

Theoretical calculations of the wave drag at super-
sonic speeds of nonlifting wings of double wedge and
biconvex section are reviewed, and the best method
of presenting the results considered. Using this


method, a representative selection of the available
numerical evaluations of the theory, are presented.
These should be of value for wing drag estimation
purposes,

N-23161*
Aeronautical Research Council (Gt. Brit.)
ASYMMETRIC TAILPLANE: LOADS DUE TO
SIDESLIP. Winfried Braun. 1953. 45p. diagrs.,
2 tabs. (ARC CP 119)

A method is derived which estimates the tailplane
rolling-moment coefficient due to sidealip for use in
strength calculations. The investigation covers the
contributions to the rolling moment from the end-
plate effect at the fin (twin fins are not considered),
the dthedral of the tailplane, the effect of the body
(which differs on the lee and windward-sides), the
effects of sweepback and plan form, and of unsym-
metrical lift distribution on the main wing. An
allowance is made for the influence of propeller
slipstream, and a tolerance suggested to cover
inaccuracies of the method. Comparison with exper-
iment shows good agreement. The method is
summarized and an example given in appendices.


N-23162*

Aeronautical Research Council (Gt. Brit.)
HIGHER HARMONICS OF FLAPPING ON THE
HELICOPTIER ROTOR. W. Stewart. 1953. 30p.
diagrs. (ARC CP 121)

The amplitudes of the flapping harmonics up to the
sixth harmonic have been calculated for a range of
blade pitch and blade inertia number for tip speed
ratios up to 0. 6. The calculations are for a
straight (infinitely rigid) blade, and blade stalling is
not taken into account. The-results are given in
the form of generalized curves. For the ordinary
helicopter, the amplitude of any harmonic is about
one-twelfth of the preceding harmonic, for a tip
speed ratio about 0. 3.


N-23174 *

Forest Products Research Lab. (Gt. Brit. )
TRIALS OF TIMBERS FOR PLYWOOD MANUJ-
FACTURE. SPECIAL TESTS, BLACK POPLAR -
POPULUS NIGRA HOME GROWN. (28 POUNDS
PER CUBIC FOOT AT 15 PER CENT MOISTURE
CONTENT) PROGRESS REPORT EIGHTEEN.
February 1953. 14p. 6tabs. (Forest Products
Research Lab. )

The surface of poplar veneer often presents a rough
or woolly appearance, particularly as it leaves the
peeler. The main object of the tests described in
this report was to see how far this surface condition
could be influenced by adjustments of the knife and
nosebar positions, by the treatment of the billets and
by preventing accumulation of "fuzz. As a
secondary object, consideration was given to the
possibility of a wider use of poplar in plywood
manufacture.






NACA
RESEARCH ABSTRACTS NOAC4


N-23198 *

Royal Aircraft Establishment (Gt. Brit.)
RESIDUAL LATTICE STRAINS IN PLASTICALLY
DEFORMED ALUMINIUM. Catherine Md. Bateman.
January 1953. 19p. diagrs., 10 tabs. (RAE Met.72)

The residual lattice strains in three types of
aluminum have been measured using X-ray diffrac-
tion methods and have been compared quantitatively
with the theory of an intergr~anar microstress
system first put forward by G. B. Greenough. A.
discrepancy between theory and experiment whn the
specimens-are in the form of tensile test pieces has
again been observed. It is concluded that these re-
sults raise doubts as to the validity of the theory in
its present form.


N-23199*

Royal Aircraft Establishment (Gt. Brit.)
ALLOYS OF TITANIUM W1TH SILICON. D. A.
Suteliffe and A. C.Spickett. February 1953. 33p.
diagrs., photos., 8 tabs. (RAE Met.73)

Binary titanium alloys containing up to 5.5-percent
silicon by weight have been prepared by are melting,
and the limit of solid solubility of silicon in titanu
determined by quenching and subsequent microscopic
examination of the specimens. Tensile and hardness
tests have been made on the alloys at temperatures
up to 6500 C and 800o C, respectively. The results
show that silicon has a useful strengthening effect on
titanium at elevated temperatures; 4.0-percent
silicon increases the U.T.S by approximately 3 times
at 6500 C. Silicon also improves the oxidation
resistance at 8000 C and 10000 C. Limited experi-
ments have been made to see whether the alloys were
heat treatable. Some observations on the working
and machining of the alloys are given.


MISCELLANEOUS


NACA Rept. 1087

Errata No. 1 on 'INTERNAL-LIQUID-FILM-
COOLING EXPERIMENTS WITH AIR-STREAM
TEMPERAURES TO 20000 F IN 2- AND 4-INCH-
DIAMETER HORIZONTAL TUBES.' George R.
Kiney, Andre~w E. Abramson and John L. 81oop.
1952.


NACA TN 2792

Errata No. 1 on "DIRECT-RIEADING DESIGN
CHARTS FOR 2r48-T3 ALUMINUM-ALLOY FLAT
COMPRESSION PANELS HAVING LONGIUDINAL
FORMED HAT-SECTION STIFFENERS AND COM-
PAISN WITHi PANELS HAVING Z-SECTIO
STIFFENERS. William A. Hickmnan and Norris F".
Dow. March 1953.


N-23183 *

Explosives Research and Development Establishment
(Gt. Brit.) THE TH~ERMIODYNAMIC PROPERTIES
OF SOME METALS AN THEIR OXIDES AT HIGH
TEMPERATURES. N. W. Luft. September 1948.
18p. 10 tabs. (ERDE Tech. Memo. 14/M1/48)

This memorandum is an attempt to determine the
thermodynamic Broperties of certain metals and
their oxides at high temperatures. The data are
required in the assessment of the value of these
metals as fuels for rocket propulsion. The metals
studied were those of low atomic weight, such as;
lithium, beryllium, boron, magnesium, aluminum,
and calcium


N-23196*

Royal Aircraft Establishment (Gt. Brit.)
A SATURABLE-CORE CURRENT REFERENCE
CIRCUIT FOR USE WITH MAGNETIC AMPLIFIERS-
A. G. Milnes and T. V. Ve~rnon. October 1952. 14p.
diagrs., tab. (RAE Tech. Note EL. 46)

A reliable, robust saturable-core type reference
circuit suitable for use with magnetic amplifier
regulators is described and its theory of operation
discussed. Details are given of a practical design
for 115V, 400 c/s operation writh an output current of
about 8 mA in a load of 60 ohms. With a +10 percent
supply voltage and +7-1/2 percent frequency change
an accuracy of better than 11l percent may be ex-
pected during 1,000 hours for a range of ambient
temperature from +150 C to +350 C. If the ambient
temperature range is wide, for exaple the service
range -65o to +700 C, rough temperature control to
within +50 C is required on two of the components
to obtain the same over-all accuracy in the refer-
ence current. In the design described the reference
draws a current of up to 50 mA from the 115 volt
supply, excluding the oven heater load; and the over-
all size of the reference circuit complete with the
temperature-control oven is 3 in. x 2-1/2 in. x



N-23197*

Royal Aircraft Establishment (Gt. Brit.)
SOME FURTHER OBSERVATIONS ON THE FATIGUE
PROCESS IN PURE ALUMINIUM. P. J. E. Forsyth.
December 1952. 25p. photos., diagrs. (RAE Met.70)

The fatigue process in pure aluminum has been more
fully studied; with special reference to the mecha-
nism of slip under cyclic stresses. The similarities
and differences of deformation produced by cyclic and
static stressing have been investigated both micro-
scopically, and with the multiple beam interference
technique. The effect of cyclic stressing at a higher
frequency has been investigated, and it is concluded
that cyclic stressing may produce a considerable
rise in temperature in the region of slip bands which
become more localized at higher frequencies.






NACA
RESEARCH ABSTRACTS NO.44


DECLASSIFIED NYACA REPORTS

NACA RM A8B05

EXPERIMENTAL INVESTIGATION OF THE
EFFECTS OF SUPPORT INTERFERENCE ON THE
DRAG OF BODIES O1F REVOLUTION AT A MACH
NUMBER OF 1.5. Edward W. Perkins. May 7, 1948.
50p. diagrs., photos. (NACA RM ABB05)
(Declasslfled from Confidential, 4/10/53)

This report contains the results of extperimental
invest igat ion to determine the effects of support
Interference on the base drag and fare drag of two
bodies of revolution. The results show that the
rea r supports affect the drag of a body of revolution
through the immediate influence on the pressures
acting over the rear portions of the body and thus
depend on the afterbody shape of the model, the
Reynolds number of flow, and the condition of the
boundary layer.

NACA RM A52A10

EFFECTS OF FINITE SPAN ON THE SECTION
CHARACTERISTICS OF TWO 450 SWEPT-BACK
WINGS: OF ASPECT RATIO 6. Lynn W. Hunton.
1March 17, 1952. 34p. diagrs. (NACA RM A52A0)
(Declassified from Restricted, 4/10/53)

A study has been made of the finite-span effects on
the local loading characteristics of two sweptback
wings at low speed with a viewv toward providing
some insight into the usefulness of two-dimensional
section data and span-loading theory for determining
the section characteristics of a swept wing. The two
wings considered were identical in plan form having
450 of sweepback of the quarter-chord line, an aspect
ratio of 6, and a taper ratio of 0. 5 but differed in
twist and in sections, the latter being the NACA
64A010 and NACA 64A810. The analysis is based on
comparisons of local pressure distributions and local
lift characteristics on the wings with'comparable
two-dimensional section data, all of which were
available at large scale.

NACA RM E8G01

PERFORMANCE OF 24-INCH SUPERSONIC AXA-
FLOW COMPRESSOR INAIR. II- PERFORMnANCE
OF COMPRESSOR ROTOR AT: EQUIVALENT TIP
SPEEDS FROM 800 TO 1765 FEET PER SECOND.
Irving A. Johnsen, Linwood C. Wright and Mlelvin
J. Rfartmann. January 21, 1949. 31p. diagrs.
(NACA RM E8GO1) (D~eclassified from Restricted,
4/13/53)

Results are presented for an experimental supersonic
axial-flow compressor over a range of equivalent tip
speeds from 800 to 1765 feet per second. The
necessity to establish simple radial eqiibrium
behind the rotor-passage-contained normal shock is
pointed out. The performance is compared to that
of a similar supersonic compressor previously run
in Freon-12 at the NACA Langley Laboratory. The
maximum pressure ratio produced in the single stage


without inlet-guide vanes was 2. 08 at 1765 feet per
second and was obtained at an adiabatic efficiency of
0. 79 and a weight flow of 64. 07 pounds per second.

NACA RM E9L21

APPROIAE RELATIVE-TOTAL -PRESSURE
LOSSES OF AN INTFINT CASCADE OF SUPER-
SONIC BLADES WITH FINITE LEADING-EDGE
THICKNESS. John F. Klapproth. March 3, 1950.
8p. diagrs. (NACA RM E9L21) (Declassified from
Restricted, 4/13/53)

By applying a hyperbolic approximation to the form
of the bow waves caused by blunt leading edges on an
infinite cascade of supersonic blades, the appro~xi-
mate losses in relative total pressure due to the
external bow-wave system, arising from blunt edges
and subsonic axial entrance velocities were com-
pared. The losses increase linearly with leading-
edge radius for any given relative Mach number. For
a relative Mach number of 1. 60, leading-edge radii
may be approximately 1. 5 percent of the normal
blade gap with a 1-percent loss of relative total
pressure.

NACA RM E51HOT

DISCUSSION OF BOUNDARY-LAYER CHARACTER-
ISTICS NEAR THE CASING OF AN AXIAL-F"LOW
COMPRESSOR. Artur Mager, John J. Mahoney and
Ray E. Budinger. December 12, 1951. 66p.
diagrs., photo., tab. (NACA RME51HO07. Now
Rept. 1085) (Declassified from Restricted, 4/10/53)

Boundary-layer velocity profiles on the casing of an
axial-flow compressor behind the guide vanes and
rotor were measured and resolved into two compo-
nents: along the streamline of the flow and perpen-
dicular to it. Boundary-layer thickness and the de-
flection of the boundary layer at the wall were the
generalizing parameters. By use of these results
and the momentum-integral equations, the character-
istics of boundary layer on the walls of axial-flow
compressor are qualitatively discussed. Important
parameters concerning secondary flow in the bound-
ary layer appear to be turning of the flow and the
product of boundary-layer thickness and streamline
curvature outside the boundary layer. Two types of
separation are shown to be possible in three dimen-
sional boundary layer.


NACA RM L8H13

THEORETICAL AND EXPERIMENTAL ANALYSIS OF
LOW-DRAG SUPERSONIC INLETS HAVING A
CIRCULAR CROSS SECTION AN A CENTRAL
BODY AT MrACH NUMBERS OF 3. 30, 2. 75, AND
2. 45. Antonio Ferri and Louis M. Nucci.
November 10, 1948. 89p. photos., diagrs. (NACA
RM L8H13) (Declassified from Confidential, 4/13/53)

Contains theoretical and experimental analysis of
circular inlets having a central body at Mach num-
bers of 3. 30, 2. 75, and 2. 45. The inlets have been
designed in order to have low drag and high pressure






NACA
RESEARCH ABSTRACTS NO.44


STRUT. Kenneth L. Wadlin, John A. Ram'm and
John R. McGehee. July 20, 1950. 31p. dagrs.
photos. (NACA RM L9K14a) (Declassified from
Confidential, 2/12/53)

An investigation was made in Langley tank No. 2 to
determine the lift and drag of a hrgh-aspect-ratio
rectangular hydrofoil supported by a single strut.
The model was tested at various depths below the
water surface at speeds up to 35 feet per second
corresponding to a Reynolds nuber of 2. Oxa 106. A
maxium lift-drag ratio of 25. 4 was obtained with
the hydrofoil at a depth of 1/2 chord. This ratio
decreased and the lift coefficient at which it occurred
increased with depth. The effects of the water sur-
face were negligible at a depth of 2 chords or
greater. Trhe data in the speed range covered
showed good agreement with corresponding aerody-
namic data from wind-tunnel tests.


NACA RM L50F21a

EXPERIMENTAL INVESTIGATION OF THE MIXIG
LOSS BEHIND THE TRAILING EDGE OF A
CASCADE OF THREE 900 SUPERSONIC TURNING
PASSAGES. Luke L. Liccint. August 15, 1950. 31p.
diagrs., photos., tab. (NACA. RM L50F21a)
(Declassified from Confidential, 4/13/53)

A two-dimensional cascade of three 90o turnin
passages, designed by the method of characteristics,
was tested at an inlet Mach number of 1. 18. The
total loss across the cascade was 9 percent. Of this
total, approximately 5 percent is chargeable to the
mixing process downstream of the trailing edges.
Although a variation of the boundary-layer thickness
occurred across the span and resulted in a nonuni-
form turning angle, the effective turning was of the
order of 900. The wake survey indicates that the
velocity tends to become unfrm very rapidly behind
the trailing edge in the blade wake.

NACA RM L50F27a

INVESTIGATION OF AN IMPULSE AXIAL-FLOW
COMPRESSOR ROTOR OVER A RANGE OF BLADE
ANGLES. Wallace MV. Schulse, John R. Erwin and
Willard R. Westphal. August 29, 1950. 34p.
diagrs., photos. (NACA RM L50F27a) (D~eclassified
from Confidential, 4/13/53)

An investigation was made of an impulse axial-flow
rotor with blades designed for 75o turning to study
its performance characteristics over a wide range of
blade-angle settings and quantity rates of flow.
Special emphasis was given to the stall-point area.
A total-pressure-rise coefficient of 3. 00 ~at 93 per-
cent efficiency, and at a different blade-angle setting
a static-pressure-rise coefficient of 0. 53 at 90 per-
cent efficiency were realized. Information on the
stall characteristics of highl cambered blading and
the conditions under which a rotor induces a flow
were also discovered.


SO


recovery. The pressure recoveries obtained are of
the same order of magnitude as those previously
obtained by inlets having very large external drag.

NACA RM L9G06

PRELIMINARY ANALYSIS OF AXIAL-FLOW
COMPRESSORS HAVING SUPERSONIC VELOCITY
AT THE ENTRANCE OF THE STATOR. Antonio
Ferri. September 12, 1949. 36p. diagrs. (NACA
RM L9G06) (Declassified from Confidential, 4/13/53)

An analysis of the phenomena related to an axial
supersonic compressor having supersonic velocity in
front of the stator is presented. The analysis shows
that comp~ression ratio for stages between 6 and 10
and adiabatic efficiency between 70 and 80 percent
can be obtained. Rotor and stator design are dis-
cussed on the basis of the results of supersonic
cascade tests. Range of performance of the com-
pressor, flow stability, and starting conditions are
also discussed.

NACA RM L9G07

ANALYTICAL AND EXPERIENTAL INVESTIGA-
TION OF 900 SUPERSONIC TURNING PASSAGES
SUITABLE FOR SUPERSONIC COMPRESSORS OR
TURINES. Luke L. Liccini. September 12, 1949.
91p. photos., diagrs., 3 tabs. (NACA RM L9G07)
(Declassified from Confidential, 4/13/53)

Four 900 two-dimensional turning passages designed
by the method of characteristics were tested at an
inlet Mach number of 1. 71. The measured losses
varied from 5 to 15 percent of the inlet stagnation
pressure. The smallest loss was obtained for a
passage in which separation ep1 the convex surface
was minimized through the introduction of a favor-
able pressure gradient-


NACA RM L9JO5a

INVESTIGATION OF AN IMPULSE AXIAL-FLOW
COMPRESSOR. John R. Erwin and Wallace M.
Schulze. February 8, 1950. 51p. diagrs., photos.
(NACA RM L9JO5a) (Declassified from Restricted,
4/13/53)

An investigation ~was made of the possibility of ob-
taining high pressure coefficients by applying the
constant-pressure principle to axial-flow compres-
sors. Use of blade sections developed by cascade
tests permitted the design of a rotor for 750 turning
and a total-pressure coefficient of 2. 40. This rotor
was tested at low speeds with and without stator
blades. A rotor total-pressure coefficient of 2. 30
with an efficiency of 98. 3 percent and a stage pres.
sure coefficient of 2. 05 with an efficiency of 89. 6
percent were measured at a design quantity-flow
coefficient of 0. 82.

NACA RM LOK14a

TANK TESTS AT SUBCAVITATIONV SPEEDS OF AlN
ASPECT-RATIO-10 HYDROFORL WITH A SINGLE






NACA II
RESEARCH ABSTRACTS NO.44

NACA RM L50G05

INV ESTIGATION OF A SH ROUDED AN AN
UNSHROUDED AX1AL-FLOW SUPERSONIC COM-
PRESSOR. Emannel Boxer and John R. Erwrin.
September 15, 1950. 54p. diagrs., photos., tab.
(NACA RM L50G05) (Declassified from Confidential,
4 13 53)

Details of the design and results of tests of an
unshrouded NACA supersonic axial-flow compressor
in Freon 12 are presented. Blade sections thicker
than previously used permitted operation without an
attached shroud. Tests were made at tip speeds
ranging from 866 feet per second to 1680 feet per
second when converted to sea-level air. At equiva-
lent tip speed of 1610 feet per second with guide
vanes, the rotor produced a pressure ratio of 2. 03
with 83. 5 percent at an air-flow rate of 28.0O pounds
per second; without vanes, the values are 2. 20, 84. 5
percent, and 27. 8, respectively. In addition,
shrouded-rator results are compared with those
obtained in air at the NACA Lewis Flight Propulsion
Laboratory


NACA RM L52B06

APPLICATION OF SUPERSONIC VORTEX-FLOW
THEORY TO THE DESIGN OF SUPERSONIC
IMPULSE COMPRESSOR- OR TUTRBINE-BLADE
SECTIONS. Emanuel Boxer, James R. Sterrett and
John Wlodaraki. April 24, 1952. 70~p. diagrs.,
photos., 4 tabs. (NACA RM L52B06) (Declassified
from Confadential, 4/13/53)

A method of designing related supersonic turbine or
compressor blade sections is developed based upon
a known solution of supersonic vortex flow. A major
part of the turrung is accomplished by concentric
circular streamlines which are generated by uniquely
shaped transition area near the leading edge of blade
surfaces. An examination of the supersonic starting
criteria was made as well as means of thickening the
leading- and trailing-edge portions. Cascade test
results of four blade passages are presented.

NJACA RM L52J27a

A SIMPLIFIED MATHEMATICAL MODEL FO
CALCULATING AERODYNAMIC LOADING AND
DOWNWASH FOR MIDING WING-FUSELAGE
COMBINATIONS WITH WINGS O]F ARBITRARY PLAN
FORlM. Martin ZIotnick and Samuel W. Robinson,
Jr. January 16, 1958. 36p. diagrs. (NACA
RM L52J27a) (Declassified from Restricted, 4/10/53)

It is shown that a satisfactory mathematical model
for a midwing wing-fuselage combination in subsonic
flow can be developed from a simple vortex-image
system. Usmng this scheme, a method for calculating
the lift on the fuselage in the presence of the wing is
presented and illustrated by a numerical examle. In
addition it is shown how the simplified mathematical
model can be used for calculating the downwash
behin the `wing: and for calculating the spanwise lift
distribution on the wing in the presence of the
fuselage.


NlACA-Langley 6-13-53 -4000




UNIVERITY OFFLORID




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INGEST IEID EV2MQR9RM_XS8V1H INGEST_TIME 2012-03-02T21:20:09Z PACKAGE AA00009235_00004
AGREEMENT_INFO ACCOUNT UF PROJECT UFDC
FILES