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 )
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:
AA00009235:00028

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National Advisory Committee for Aeronautics



Research Abstracts


NO.66

CURRENT NACA REPORTS

NACARept. 1140

CHARTS AND APPROXIMATE FORMULAS FOR THE
ESTIMATION OF AEROELASTIC EFFECTS ON THE
LOADING OF SWEPT AND UNSWEPT WINGS.
Franklin W. Diederich and Kenneth A. Foss. 1953.
ill, 48p. diagrs., 3 tabs. (NACA Rept. 1140.
Formerly TN 2608)

Charts are presented for the estimation of aero-
elastic effects on spanwise lift distribution, lift-
curve slope, aerodynamic center, and damping in
roll of swept and unswept wings at subsonic and
supersonic speeds. Some design considerations
brought out by the results of this paper are dis-
cussed.




NACA TM 1375

ON THE THREE-DIMENSIONAL INSTABILITY OF
LAMINAR BOUNDARY LAYERS ON CONCAVE
WALLS. (Uber eine dreidimensionale Instabilitat
laminarer Grenzschichten an konkaven Wiinden). H.
G6rtler. June 1954. 32p. diagrs. (NACA TM 1375.
Trans. from Gesellschaft der Wissenschaften zu
Gbttingen, Nachrichten, Mathematik, v. 2, no. 1,
1940)

A study is made of the stability of laminar boundary-
layer profiles on slightly curved walls relative to
small disturbances that result from vortices whose
axes are parallel to the principal direction of flow.
The result is an eigenvalue problem by which, for a
given undisturbed flow at a prescribed wall, the am-
plification or decay is computed for each Reynolds
number and each vortex thickness. For neutral dis-
turbances (zero amplification) a critical Reynolds
number is determined for each vortex distribution.
The numerical calculation produces amplified dis-
turbances on concave walls only.



NACA RM E54E03

EFFECT OF HIGH PRESSURE ON SMOKING TEN-
DENCY OF DIFFUSION FLAMES. Glen E. McDonald
and Rose L. Schalla. June 1954. 7p. diagrs., tab.
(NACA RM E54E03)

The variation in smoking tendency of a diffusion
flame was measured over a pressure range of 4 to
20 atmospheres (abs) for the two fuel types ethane and
ethylene. The fuels were burned as diffusion flames
inside a pressure vessel equipped with windows for
observation of the flame. It was found that the


*AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 15t12 H ST., NW.,
THE REPORT TITLE AND AUTHOR.
'2?If30Oi


JULY 13, 1954

changes in flame height. which measures the relative
amount of fuel that can be burned withoutt producing
smo.'ke, %ary inversely wiutl..-ieprssurP.-


NACA TN 3130

A PROCEDURE FOR THE DESIGN OF AIR-HEATED
ICE-PREVENTION SYSTEMS. Carr B. Neel, Jr.
June 1954. 63p. diagrs., photo. (NACA TN 3130)


The procedure to be followed in the design of air-
craft ice-prevention equipment in which the compo-
nents are protected by means of internally circulated
heated air is outlined. In addition to presentation of
the required heat-transfer and air-pressure-loss
equations, a simple electrical analogue is described
which was devised to facilitate the desigrt oT an air.-
heated system. An illustration is given of the
application of the analogue to a design 'prqdlem.


NACA TN 3153 JUL 15 l

VARIATION OF LOCAL LIQUID-WATER CONCEN-
TRATION ABOUT AN ELLIPSOID OF FINENESS
RATIO 5 MOVING IN A DROPLET F I E D. Rberm- --
G. Dorsch and Rinaldo J. Brun. July 1954'. 6tp. t .
diagrs., photos., 2 tabs. (NACA TN 3153) --


~. 1
f <
/
/
/


Analyses of calculated water-droplet trajectories
show that the concentration of liquid water at various
points about an ellipsoid of revolution moving through
a droplet field varies considerably. Curves of local
concentration factor as a function of spatial position
are presented in terms of dimensionless parameters.


NACA TN 3154

INFRARED SPECTRA OF 47 DICYCLIC HYDRO-
CARBONS. John H. Lamneck, Jr., Harold F.
Hipsher and Virginia 0. Fenn. June 1954. 34p.
diagrs., 5 tabs. (NACA TN 3154)

The infrared spectra are presented for 47 dicyclic
hydrocarbons consisting of some alkyldiphenyl-
methanes, alkyldicyclohexylmethanes, alkylnaphtha-
lenes, alkyltetralins, 1,3-diphenyl-2-alkylpropanes
and 1,3-dicyclohexyl-2-alkylpropanes. The physical
properties of these highly purified compounds arae.
included for reference purposes. .

NACA TN 3161 /

AN INVESTIGATION OF THE U E 9F ifq 1
POWERED MODELS FOR GUST LOAd IDlES
WITH AN APPLICATION TO A AILL.ESS SWEPT-
WING MODEL AT TRANSONIC SPEEW. A. James ,
Vitale. H. Press and C. C. Shuflfargbe., June. -'
1954. 36p. diagrs., phointos.. tab. (ACA. q1J16I..


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






NACA
RESEARCH ABSTRACTS NO.66


Flight tests in continuous rough air at transonic
speeds have been made of a tailless 450 sweptback-
wing rocket model as part of a program to investi-
gate the feasibility of using existing rocket-model
techniques for obtaining gust-loads information. The
results of airplane flight surveys for determining a
working basis for forecasting and selecting suitable
test days are presented. Use is made of power-
spectral methods of generalized harmonic analysis
in presenting both the experimental and calculated
characteristics of load response in rough air of a
configuration having low damping in pitch.




NACA TN 3168

A NEW HODOGRAPH FOR FREE-STREAMLINE
THEORY. Anatol Roshko, California Institute of
Technology. July 1954. 39p. diagrs., 3 tabs.
(NACA TN 3168)

The Helmholtz-Kirchhoff method for separated flow
past a flat plate normal to the stream is modified to
allow arbitrary separation velocity and base pres-
sure so that values more in conformity with experi-
ment may be chosen. The solution depends on a
single base-pressure parameter k. The computa-
tions depend on a particular choice of free-
streamline hodograph which gives a definite wake
width for every value of k. In this way the wake
width is correlated with the drag. Examples are
worked out for a wedge and a circular cylinder.





NACA TN 3169

ON THE DRAG AND SHEDDING FREQUENCY OF
TWO-DIMENSIONAL BLUFF BODIES. Anatol
Roshko, California Institute of Technology.
July 1954. 29p. diagrs., tab. (NACA TN 3169)

A semiempirical study was made of the bluff-body
problem. The relation between the wake and the
potential flow outside the wake and cylinder was
studied and some experiments were made with inter-
ference elements in the wake close behind a cylinder.
Free-streamline theory was combined with experi-
mental results to obtain a correlation between bluff
cylinders of different shapes.




NACA TN 3171

SOME NEW DRAG DATA ON THE NACA RM-10
MISSILE AND A CORRELATION OF THE EXISTING
DRAG MEASUREMENTS AT M = 1.6 AND 3.0.
Robert J. Carros and Carlton S. James. June 1954.
24p. diagrs., photos., tab. (NACA TN 3171)
Measurements of the zero-lift total drag of the NACA
RM-10 missile were made on gun-launched, free-
flight models at Mach numbers of 1.6 and 3.0 and
corresponding Reynolds numbers of 3.0 million and
5.0 million. Results showed transition location to
have an important effect on the drag. Results of
this and several other investigations were correlated
on the basis of considerations of Mach number,
Reynolds number, transition location, and heat-
transfer effects.


NACA TN 3180

DETERMINATION OF VISCOSITY OF EXHAUST-
GAS MIXTURES AT ELEVATED TEMPERATURES.
J. C. Westmoreland, National Bureau of Standards.
June 1954. 41p. diagrs., photo., 10 tabs. (NACA
TN 3180)

The viscosities of five samples of dry-exhaust-gas
mixtures at approximately atmospheric pressure and
at temperatures of from 00 to 1,1000 C are presented
A method of analysis and correlation proposed by
Sutherland has been applied to the data and equations
were derived therefrom for representing the results.
It is believed that these equations may be used for
computing viscosities of all mixtures having compo-
sitions within the range of the five studied at any
temperature up to about 1,3000 C. The effect of
moisture upon the viscosity of exhaust gas was com-
puted by a relation developed from experimental
values for moist air.




NACA TN 3185

TABLES FOR THE COMPUTATION OF WAVE DRAG
OF ARROW WINGS OF ARBITRARY AIRFOIL SEC-
TION. Frederick C. Grant and Morton Cooper.
June 1954. 9p. diagr., 28p. tabs. (in pocket) (NACA
TN 3185)

Tables and computing instructions are presented for
the rapid evaluation of the wave drag of delta wings
and of arrow wings having a ratio of the tangent of
the trailing-edge sweep angle to the tangent of the
leading-edge sweep angle from -1.0 to 0.8. The
tables cover a range of both subsonic and supersonic
leading edges.






NACA TN 3187

THE NEAR NOISE FIELD OF STATIC JETS AND
SOME MODEL STUDIES OF DEVICES FOR NOISE
REDUCTION. Leslie W. Lassiter and Harvey H.
Hubbard. July 1954. 38p. diagrs., photos. (NACA
TN 3187)

Experimental studies of the pressure fluctuations
near jets were made during unchoked operation of
both a full-scale turbojet engine and a 1-inch-
diameter high-temperature model jet and during
choked operation of model jets of 0.275-inch to 2.00-
inch diameter with unheated air. Frequency spectra
and spatial distributions of pressure magnitude are
given for the full-scale configuration and model data
are used to illustrate probable trends at different
operating conditions for the unchoked jets. Model
tests for choked operation indicate the presence of a
discrete-frequency component, and shadowgraph
records illustrate that an unusual type of flow forma-
tion is associated with this condition. The frequency
of this component is shown to be somewhat related to
the shock-separation distances and to nozzle
diameter. Laboratory methods of reducing the
magnitude of pressure fluctuations from choked and
unchoked jets are discussed and sample illustrations
are given.







NACA
RESEARCH ABSTRACTS NO. 66

NACA TN 3188

AN ANALYTICAL INVESTIGATION OF AIRPLANE
SPIN-RECOVERY MOTION BY USE OF ROTARY-
BALANCE AERODYNAMIC DATA. Stanley H. Scher.
June 1954. 38p. diagrs., tab. (NACA TN 3188)

A method for analytically investigating the motions
of an airplane while in a spin or a recovery from the
spin by use of applicable equations of motion,
modified wind-tunnel rotary-balance measurements,
and spin-geometry relationships is presented and
applied to the step-by-step calculation of the spin
recovery for one airplane configuration. Difficulties
encountered in applying the rotary-balance data in
the calculations are discussed, and it is pointed out
that certain inconsistencies must be cleared up
before the method can be accepted as adequate to
give detailed spin-recovery motions for a specific
airplane.


NACA TN 3193

AN EXPLORATORY INVESTIGATION OF SKIN
FRICTION AND TRANSITION ON THREE BODIES
OF REVOLUTION AT A MACH NUMBER OF 1.61.
John H. Hilton, Jr. and K. R. Czarnecki. June
1954. 15p. diagrs. (NACA TN 3193)

The present report contains results of an explora-
tory investigation of the effects of pressure gradient
on skin friction and boundary-layer transition of
three bodies of revolution at a Mach number of 1.61.
The pressure gradients investigated were those ob-
tained on an ogive-cylinder, a cone-cylinder, and a
blunt-base (NACA RM-10) parabolic body, all having
a fineness ratio of 12.2. Tests were made at zero
angle of attack and over a Reynolds number range,
based on body length, from about 2.5 x 106 to
37 x 106.



NACA TN 3194

STATISTICAL MEASUREMENTS OF CONTACT
CONDITIONS OF 478 TRANSPORT-AIRPLANE
LANDINGS DURING ROUTINE DAYTIME OPERA-
TIONS. Norman S. Silsby. June 1954. 32p. diagrs.,
photos., 3 tabs. (NACA TN 3194)

Statistical measurements have been obtained from
photographs taken with a specially built motion-
picture camera of 478 landings of present-day trans-
port airplanes during routine daylight operations in
clear air at the Washington National Airport. From
these measurements, sinking speeds, bank angles,
rolling velocities, and horizontal speeds have been
evaluated and a limited statistical analysis of the re-
sults has been made. An attempt was made to deter-
mine the effect of various parameters such as gusty-
wind conditions, wing loading, and size of airplane
(number of engines) which influence the landing con-
tact conditions.



NACA TN 3195

TIME-TEMPERATURE PARAMETERS AND AN
APPLICATION TO RUPTURE AND CREEP OF
ALUMINUM ALLOYS. George J. Heimerl. June
1954. 35p. diagrs., tab. (NACA TN 3195)


3

The application of time-temperature parameters to
stress-strain, rupture, and creep data for metals and
alloys is reviewed. Some comparisons are made of
theoretical and experimental parameters. A param-
eter based upon rate-process theory was successfully
applied to rupture and creep data for aluminum and
various aluminum alloys. The value of the constant
in the parameter, which provided the best correla-
tion of the data. was determined for each material
and application. Master curves of stress against the
parameter which summarize extensive data on the
aluminum alloys are presented for rupture, minimum
creep rate, and time to 1 or 2 percent strain. Pre-
dictions of long-time life from short-time data are
shown to be possible.



NACA TN 3198

DYNAMIC STABILITY AND CONTROL CHARACTER-
ISTICS OF A CASCADE-WING VERTICALLY RISING
AIRPLANE MODEL IN TAKE-OFFS, L -%DL\GS.
AND HOVERING FLIGHT. Marion 0. McKinney,
Louis P. Tosti and Edwin E. Davenport. June 1954.
45p. diagrs., photos., tab. (NACA TN 3198)

An investigation has been made to determine the sta-
iiir.m and control characteristics of a cascade
vertically rising airplane model in the take-off,
landing, and hovering phases of flight. The model
had four propellers with the shaft axes essentially
parallel to the fuselage axis and distributed along
the span so that the four wings, which were arranged
in cascade relation to turn the slipstream downward
approximately 900, were completely immersed in
the slipstream of the propellers.




NACA TN 3202

OSCILLATING PRESSURES NEAR A STATIC
PUSHER PROPELLER AT TIP MACH NUMBERS UP
TO 1.20 WITH SPECIAL REFERENCE TO THE EF-
FECTS OF THE PRESENCE OF THE WING. Harvey
H. Hubbard and Leslie W. Lassiter. July 1954.
35p. diagrs., photos., tab. (NACA TN 3202)

Static tests were conducted for the measurement of
free-space usrllaiirg pressures near the propeller
in the region where a wing might be located for the
tip Mach number range of 0.50 to 1.20. Some
measurements were also made near the tips in the
region of the fuselage in order to extend the range of
existing work from a tip Mach number of 1.00 to
1.20. Some comparisons of these results with theory
are given. In addition to a description of the free-
space pressure field, measurements are given for
some of the same field points with a wing in place.
Charts based on experimental data are presented to
enable a designer to estimate the maximum free-
space oscillating pressures near the propeller for
tip Mach numbers up to 1.20.




NACA TN 3204

AN INVESTIGATION OF THE CREEP LIFETIME OF
75S-T6 ALUMINUM-ALLOY COLUMNS. Eldon E.
Mathauser and William A. Brooks, Jr. July 1954.
28p. diagrs., photos., tab. (NACA TN 3204)






4


The results of short-time elevated-temperature
creep tests of 75S-T6 aluminum-alloy columns are
presented and examined with the objective of obtain-
ing procedures for predicting column lifetime.
Semiempirical lifetime curves are obtained with the
aid of a previously published column creep theory
and are used for deriving column curves. A study is
made of the effects of variations of stress and out-of-
straightness on column lifetime. Small variations in
out-of-straightness have been found to be of little
practical significance; whereas, small stress vari-
ations change the column lifetime considerably.
Plots that do not explicitly include out-of-straightness
are presented and may be satisfactory for predicting
column lifetime for design purposes.




NACA TN 3214

FUNDAMENTAL STUDY OF EROSION CAUSED BY
STEEP PRESSURE WAVES. B. G. Rightmire and
J. M. Bonneville, Massachusetts Institute of
Technology. June 1954. 30p. diagrs., photos.,
2 tabs. (NACA TN 3214)

A fundamental study of erosion caused by steep pres-
sure waves has been carried out. It is believed that
the study gives an insight to the possible causes of
damage in high-speed sleeve bearings. In particular,
the effect on annealed copper surfaces of steep-
fronted pressure waves in oil has been studied, the
general conclusion being that cavitation of the oil is
the probable cause of damage.



NACA TN 3220

AERODYNAMIC LOADS ON A LEADING-EDGE
FLAP AND A LEADING-EDGE SLAT ON THE
NACA 64A010 AIRFOIL SECTION. John A. Kelly
and George B. McCullough. June 1954. 33p.
diagrs., 8 tabs. (NACA TN 3220)

Loads data in the form of chordwise distributions of
pressure, normal-force, chord-force, and moment-
coefficients are presented for a leading-edge flap and
leading-edge slat installed on a two-dimensional
NACA 64A010 airfoil. Pressure data for various
model arrangements are presented in tabular form.
The data were obtained for a Reynolds number of
6 million.






NACA TN 3221

STUDY OF THE SUBSONIC FORCES AND MOMENTS
ON AN INCLINED PLATE OF INFINITE SPAN.
Bradford H. Wick. June 1954. 25p. diagrs. (NACA
TN 3221)

A study has been made of the flow about an inclined
plate, the forces on the plate, the adequacy of theory
in predicting the forces, and the extent to which the
force characteristics of the plate are indicative of
those of thin airfoil sections.


NACA
RESEARCH ABSTRACTS NO.66



NACA TN 3222

MEASUREMENT OF HEAT TRANSFER IN THE TUR-
BULENT BOUNDARY LAYER ON A FLAT PLATE
IN SUPERSONIC FLOW AND COMPARISON WITH
SKIN-FRICTION RESULTS. C. C. Pappas. June
1954. 32p. diagrs., tab. (NACA TN 3222)

Local heat-transfer rates and average skin-friction
coefficients on the surface of a heated flat plate at
zero incidence to the air stream at Mach numbers of
1.69 and 2.27 are presented for a Reynolds number
range of 106 to 107. The variation of heat transfer
with Mach number was found to be the same as that
of directly measured skin friction on unheated bodies.


NACA TN 3228

AERODYNAMIC INVESTIGATION OF A FOUR-
BLADE PROPELLER OPERATING THROUGH AN
ANGLE-OF-ATTACK RANGE FROM 00 TO 1800.
H. Clyde McLemore and Michael D. Cannon. June
1954. 62p. diagrs., photo. (NACA TN 3228)

The results of an investigation in the Langley full-
scale tunnel of the aerodynamic characteristics of a
four-blade, 5. 33-foot diameter propeller are pre-
sented for angles of attack from 00 to 1800, blade
angles from 00 to 67. 50, and a range of advance
ratio from 0 to 6. 2. The results include a prelim-
inary exploration of vertical descent and a compar-
ison with theory of the rate of change of the normal-
force coefficient with angle of attack and the aero-
dynamic characteristics of the propeller at zero
angle of attack.


NACA TN 3230

INVESTIGATION OF DISTRIBUTED SURFACE
ROUGHNESS ON A BODY OF REVOLUTION AT A
MACH NUMBER OF 1.61. K. R. Czarnecki, Ross
B. Robinson and John H. Hilton, Jr. June 1954.
35p. diagrs., photo., 2 tabs. (NACA TN 3230)

An investigation has been conducted in the Langley
4- by 4-foot supersonic pressure tunnel to study the
effect of surface roughness on the skin friction and
temperature-recovery factor of an ogive-cylinder
body of revolution at a Mach number of 1.61. Tests
of models having four different surface roughnesses
were made at zero angle of attack through a Reynolds
number range (based on body length) of 2.5 x 106 to
37 x 106. The drag, base pressure, and surface
temperatures were measured for conditions of
natural and fixed transition.




BRITISH REPORTS



N-30870*

Aeronautical Research Council (Gt. Brit.)
A BRIEF REVIEW OF THE PROBLEM OF EXHAUST
SILENCING. F. L. West. 1954. 13p. diagrs.
(ARC R & M 2803; ARC 9714. Formerly RAE Tech.
Note Gas 20)






NACA
RESEARCH ABSTRACTS NO. 66


This note reviews past and present work on noise re-
duction of the reciprocating engine exhaust.
Collected measurements of the noise level surround-
ing various engine installations and the effect of
silencing experiments on engine and aircraft per-
formance are presented. Some observations of gas
turbine noises are also included.




N-3087f

Aeronautical Research Council (Gt. Brit.)
REPORT ON THE FLOW PHENOMENA AT SUPER-
SONIC SPEED IN THE NEIGHBOURHOOD OF THE
ENTRY OF A PROPULSIVE DUCT. G. H. Lean.
1954. 15p. diagrs., photos. (ARC R & M 2827.
Formerly ARC 11,868; EAP 94)

With a parallel entry duct followed by a straight di-
vergent diffuser of 100 total angle, the flow inside
the parallel tube was supersonic provided the outlet
pressure of the diffuser was less than a certain criti-
cal value. For pressure higher than the critical
value, the flow became subsonic at the duct entry, a
shock wave was formed at the entrance lip and the
rate of air flow through the tube decreased. For out-
let pressure less than critical, the flow was super-
sonic for a distance inside the duct entry depending
on the outlet pressure, the flow becoming subsonic
further along the duct. Annular type intakes gave
lower pressure recoveries than the unobstructed
type.




N-30872 *

Aeronautical Research Council (Gt. Brit.)
THE USE OF TENSOR NOTATION TO DEVELOP
CHARACTERISTIC EQUATIONS OF SUPERSONIC
FLOW. C. N. H. Lock and R. C. Tomlinson. 1954.
18p. diagrs. (ARC R & M 2632; ARC 11,584. For-
merly ARC 12,206; FM 1341)

General equations of steady motion of a nonviscous
fluid are given in tensor notation. It is assumed that
one family of coordinate surfaces, xa = constant, are
characteristic surfaces, that is, surfaces in which
the transverse derivatives of the flow variables are
not determined by their values on the surface itself.
The condition for this gives the result that the veloc-
ity normal to the surface is sonic. The relation
which must then hold between the variables on the
surface itself is also determined (characteristic
equation). Simplicity gives a deeper insight into
behavior of axisymmetric and two-dimensional flows.




N-30873*

Aeronautical Research Council (Gt. Brit.)
THE FLUTTER OF SWEPT AND UNSWEPT WINGS
WITH FIXED-ROOT CONDITIONS. PART I WIND-
TUNNEL EXPERIMENTS. PART II COMPARISON
OF EXPERIMENT AND THEORY. PART III WING
TORSIONAL STIFFNESS CRITERION. W. G.
Molyneux. 1954. 23p. diagrs., 4 tabs. (ARC
R & M 2796; ARC 13,306. Formerly RAE Structures
58)


5


The "fixed-root" flexure-torsion flutter characteris-
tics of four model wings of different taper ratios have
been investigated in the wind tunnel. The wing
interia axis and the angle of sweepback have been
varied on each wing over the ranges 0.4c and 0.5c and
00 to 500, respectively.


N-30874*

Aeronautical Research Council (Gt. Brit.)
CRITICAL MACH NUMBERS FOR THIN UNTAPERED
SWEPT WINGS AT ZERO INCIDENCE. S. Neumark.
1954. 74p. diagrs., 12 tabs. (ARC R & M 2821;
ARC 13,541. Formerly RAE Aero 2355)

This report is a continuation of two previous papers.
Subsonic flow is further investigated using linear
theory. Methods for calculating "lower" and "upper"
critical Mach numbers are given, the solution for the
main problem being preceded by a short analysis of
critical Mach numbers for the simpler cases of infi-
nite wings. Critical Mach numbers for wings with
four different profiles, showing the effect of thickness
ratio and of sweepback angle, are determined. The
method applies strictly to wings of large aspect ratio.
Several conclusions concerning the practical use of
swept-wing design are presented.



N-30875*

Aeronautical Research Council (Gt. Brit.)
DETERMINATION OF REVERSAL SPEED OF A
WING WITH A PARTIAL-SPAN FLAP AND INSET
AILERON. W. G. Molyneux and E. G. Broadbent.
1954. 0lp. diagrs., tab. (ARC R & M 2793; ARC
13,308. Formerly RAE Tech. Note Structures 23)

Control reversal due to deformation of a wing with a
partial-span flap and inset aileron is considered
theoretically for the particular case of a flap held at
the root end. The semirigid method is used. An
investigation is made for a particular aircraft. The
calculated reversal speed is found to be considerably
lower than for the straight-forward wing-aileron
case. The effect of variation of the degrees of wing
and flap constraint is also considered. It is con-
cluded that an increase in reversal speed is best ob-
tained by an increase in flap root stiffness.



N-30876*

Aeronautical Research Council (Gt. Brit.)
WIND-TUNNEL TESTS ON TWO-DIMENSIONAL
SUPERSONIC AEROFOILS AT M = 1.86 AND
M = 2.48. D. Beastall and R. J. Pallant. 1954.
20p. diagrs., photos., 4 tabs. (ARC R & M 2800;
ARC 13,768. Formerly RAE Aero 2384)

Unswept double-wedge and circular-arc airfoils were
used to study the viscous effects which are not
accounted for in the linearized and shock-expansion
airfoil theories. Characteristics derived from
measured surface pressures are compared with
theoretical values. Schlieren observation was em-
ployed to observe the flow and, in particular, the
separation near the trailing edges. Wires caused a
delay in separation; pitot-tube traverses indicated
the velocity profiles. The position of separation was
also examined by means of oil.








6


N-30877*

Aeronautical Research Council (Gt. Brit.)
GENERAL PERFORMANCE CALCULATIONS FOR
GAS TURBINE ENGINES. D. H. Mallinson. 1954.
58p. diagrs. (ARC R & M 2684; ARC 10,552. For-
merly Power Jets Rept. R. 1214)

An attempt is made to summarize theoretical work
carried out during the past few years aimed at dis-
covering the potentialities of the gas turbine as
power plant in many fields of application, especially
as an aircraft power unit. The performance of
various modification of the ideal gas turbine cycle is
considered in some detail. Works of various authors
are combined and edited in order to depict perform-
ance attainable by practical engines. Association
between the gas turbine and jet reaction is con-
sidered. A method of estimating performance of a
simple jet engine is detailed.




N-30878*

Aeronautical Research Council (Gt. Brit.)
VELOCITY DISTRIBUTION ON THIN BODIES OF
REVOLUTION AT ZERO INCIDENCE IN INCOM-
PRESSIBLE FLOW. S. Neumark. 1954. 42p.
diagrs. (ARC R & M 2814; ARC 13,756. Formerly
RAE Aero 2389)

A new method of determining velocity distribution on
slender bodies of revolution in axial flow is ex-
pounded, analogous to the linear perturbation method
widely used for slender symmetrical profiles in two
dimensions. The proposed method leads to simple
approximate formulas for velocity distribution on a
body, once the equation of the meridian line is given,
either in the form of a polynomial, of a square root
of one. The method has been used for computing
velocity distributions on 12 different bodies, of seven
different thickness ratios each, so as to exhibit the
most characteristic features in typical cases, and to
show some effects of thickness changes.





N-30990*

Royal Aircraft Establishment (Gt. Brit.)
PRESSURE DISTRIBUTIONS ILLUSTRATING FLOW
REATTACHMENT BEHIND A FORWARD MOUNTED
FLAP. E. C. Maskell. March 1954. 23p. diagrs.,
tab. (RAE Tech. Note Aero 2295)

A series of pressure distributions over anNACA.0015
wing with a forward mounted 5-percent chord split
flap are presented. The results, obtained by Clarke
in 1946, have not previously been published. Two
basic types of pressure distribution are identified;
and these are associated with fully detached flow be-
hind the flap, and with reattachment of the flow to the
wing surface. An intermediate type corresponding,
it is suggested, to reattachment in the neighborhood
of the wing trailing edge is also identified. The
physical nature of the flow behind a flap is considered
and some ideas on the mechanism of reattachment
are suggested.


NACA
RESEARCH ABSTRACTS NO.66



N-32012*

Aeronautical Research Council (Gt. Brit.)
DOWNWASH MEASUREMENTS BEHIND A 12-FT
DIAMETER HELICOPTER ROTOR IN THE 24-FT
WIND TUNNEL. R. A. Fail and R. C. W. Eyre.
1954. 12p. diagrs., 9 tabs. (ARC R & M 2810;
ARC 12,895. Formerly RAE Tech. Note Aero 2018)

Some measurements of downwash have been made in
a plane behind a 12-ft diameter helicopter rotor over
a range of shaft inclination and tip speed ratio. In
the various operation conditions, the tunnel tests are
in reasonable agreement with the theoretical results
for the appropriate type of loading.








N-32013*

Aeronautical Research Council (Gt. Brit.)
24-FT. WIND TUNNEL TESTS ON A PROPELLER
WITH NACA 16 SERIES SECTIONS. TEST RESULTS
AND ANALYSIS INTO MEAN LIFT-DRAG DATA.
A. R. C. MacDougall and A. B. Haines. 1954. 18p.
diagrs., 6 tabs. (ARC R & M 2602; ARC 11,817.
Formerly RAE Aero 2276)

The de Havilland propeller was designed to give good
performance at high speeds and the aim of the tests
was to check whether, as a result, serious losses in
take-off performance had been incurred. Results
are reassuring and have been analyzed into lift and
drag data for use with the standard single-radius
method. Data are compared with results from Clark
Y propellers. Improvement becomes greater with
increase in Mach number, and so NACA series 16
sections appear to have a beneficial effect on the
stalling performance at all tip speeds.







N-32014*

Aeronautical Research Council (Gt. Brit.)
HEAT TRANSFERENCE AND PRESSURE LOSS FOR
AIR FLOWING IN PASSAGES OF SMALL DIMEN-
SIONS. J. Remfry. 1954. iv, 56p. diagrs., photos.,
4 tabs. (ARC R & M 2638)

This investigation had as its primary object the ex-
perimental determination of the heat-transfer and
pressure-loss characteristics for air flowing in
small triangular, square, hexagonal, and round pass-
ages. The heat interchanger models, with a frontage
6 inches square, each comprised from just under 150
to over 2,250 passages, according to their size and
spacing. The hydraulic diameter of the smallest
tubes was about 0.08 inch. The heat transfer in
small smooth passages was found to be less than that
usually measured for turbulent flow in tubes of larger
size, and there was a tendency for a prolonged transi-
tion.







NACA
RESEARCH ABSTRACTS NO.66


N-32015*

Aeronautical Research Council (Gt. Brit.)
THE DERIVATION OF AIRWORTHINESS PERFORM-
ANCE CLIMB STANDARDS. F. G. R. Cook and A. K.
Weaver. 1954. 64p. diagrs., 3 tabs. (ARC
R & M 2631; ARC 12,079. Formerly AAEE/Res/239)

A more rational basis for deriving climb standards
is proposed. The conception is introduced of a
"datum" performance, below which conditions pre-
disposing to an accident exist, and the level of
safety judged by an "incident rate" which is the fre-
quency with which the operational performance of
aircraft falls below this datum. A standard is
chosen so that when the aircraft type complies, the
incident rate will not exceed some tolerable value.
The effect of engine failure is included by taking
account of the probability of engine failure and the
associated loss in performance.





N-32016*

Aeronautical Research Council (Gt. Brit.)
MEASUREMENTS OF MID-CHORD PITCHING
MOMENT DERIVATIVES AT HIGH SPEEDS. J. B.
Bratt and A. Chinneck. 1954. 36p. diagrs., photos.,
9 tabs. (ARC R & M 2680. Formerly ARC 10,709;
0.672; FM 1125)

Measurements were made on a 7-1/2 percent bicon-
vex airfoil oscillating about a midchord axis in a
high-speed wind tunnel by the method of decaying
oscillations. The effect of variation of frequency
parameter was also investigated, and conditions
giving rise to sustained or growing oscillations at
subsonic speeds were examined. Comparison with
existing flat plate theories for supersonic flow shows
complete disagreement in the trend of the damping
with Mach number change, the linearized theory for
a flat plate giving an increasing positive value. A
recent theory which takes into account the shape of
the profile agrees in trend with experiment.





N-32017*

Aeronautical Research Council (Gt. Brit.)
AN INVESTIGATION OF THE DISTURBANCES
CAUSED BY A REFLECTION PLATE IN THE
WORKING-SECTION OF A SUPERSONIC WIND TUN-
NEL. A. 0. Ormerod. 1954. 16p. diagrs. (ARC
R & M 2799; ARC 13,915. Formerly RAE Tech.
Note Aero 2084)

In the region above the plate, two main disturbances
were found; there was a small disturbance from the
leading edge and a disturbance further downstream
which had originated beneath the plate. Between the
two there was a region of approximately constant
pressure in which a model could be located. The
forward disturbance seemed unavoidable. Increasing
the area of the passage beneath the plate by a recess
in the tunnel wall was the most effective way of
moving back and reducing the magnitude of the
disturbance at the rear.


N-32018*

Aeronautical Research Council (Gt. Brit.)
DESIGN AND USE OF COUNTING ACCELEROM-
ETERS. J. Taylor. 1954. lip. diagrs., photos.,
3 tabs. (ARC R & M 2812; ARC 13,587. Formerly
RAE Structures 78)

The fundamental principles underlying acceleration
recording by means of a counting accelerometer are
analyzed. The essential design requirements for a
counting accelerometer are presented. A design
that has been specially made to meet these require-
ments is described. Both mechanical and electrical
counting are considered, but mechanical counting is
found to be superior.





N-32020*

Aeronautical Research Council (Gt. Brit.)
LIST OF CURRENT PAPERS PUBLISHED BY THE
AERONAUTICAL RESEARCH COUNCIL, NOS. 101-
150. 1954. 5p. (ARC CP 150)







N-32021*

Aeronautical Research Council (Gt. Br i.i
EXPERIMENTS AT M = 1.8 ON BODIES OF REVOLU-
TION HAVING OGIVAL HEADS. H. K. Zienkiewicz,
A. Chinneck, C. J. Berry and P. J. Peggs. 1954.
lip. diagrs., photos., 2 tabs. (ARC CP 148)

Measurements have been made in the N.P.L. 9- by
3-inch high-speed wind tunnel of the pressure distri-
butions at a Mach number of 1.8 on three bodies of
revolution with ogival heads of different shapes at
zero incidence. Comparison between the experi-
mental pressures and those obtained iricoreicails
showed good agreement.




N-32022*

Aeronautical Research Council (Gt. Brit.)
ANALYSIS OF FLIGHT MEASUREMENTS ON THE
AIRBORNE PATH DURING TAKE-OFF. W. R.
Buckingham and D. Lean. 1954. 26p. diagrs., 2 tabs.
(ARC CP 156)

Analysis of a series of systematic take-off tests with
a Meteor IV aircraft has shown that to a good approx-
imation the minimum airborne path to 50 feet may be
treated as an arc of a circle. With this assumption,
it is a simple process to derive the mean equivalent
lift coefficient used during this part of the take-off.
It has been found that the total equivalent lift coeffi-
cient used during the airborne phase decreases with
increase in the ratio of the airspeed to the stalling
speed in a simple manner, which is independent of
the thrust/weight ratio when the shortest possible
distance is required.




UNIVERSITY OF FLORIDA


3 1262 08153 066 8
8l[10i 61ll


N-32023*

Aeronautical Research Council (Gt. Brit.)
THEORETICAL LOAD DISTRIBUTIONS ON WINGS
WITH CYLINDRICAL BODIES AT THE TIPS. D. E.
Hartley. 1954. 39p. diagrs., tab. (ARC CP 147)

The effect of tip-tanks on spanwise lift distributions
is investigated theoretically for the case of minimum
induced drag in incompressible potential flow.
Charts enabling rapid estimation of the changes in
span loadings, total lifts, and induced drags are pre-
sented for a practical range of the ratio to tank
diameter to wing span. It is shown how the results
may be applied approximately to wings of any plan
form, including those with sweep or of low aspect
ratio.



N-32024*

Aeronautical Research Council (Gt. Brit.)
CRITERIA FOR CONDENSATION FREE FLOW IN
THE R.A.E.NO. 18 (9 IN. x 9 IN.) SUPERSONIC TUN-
NEL. D. J. Raney and B. Beastall. 1954. 18p.
diagrs. (ARC CP 164)

Tests were made to determine the dryness of the air
circulating through the tunnel necessary to ensure
condensation free flow through the working section.
Indications are that a strong shock wave has little
effect in causing premature condensation but that
local expansions around a model may substantially
reduce the critical humidity.



N-32025*

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE EFFECT OF THICKNESS AND
ASPECT RATIO ON THE DAMPING OF PITCHING
OSCILLATIONS OF RECTANGULAR WINGS MOVING
AT SUPERSONIC SPEEDS. W. E. A. Acum. 1954.
16p. diagrs. (ARC CP 151)

An estimate is made of the effect of variation of
aspect ratio and thickness parameter on the damping
derivative of a rectangular wing performing pitching
oscillations in a supersonic stream. After a dis-
cussion of various theories of unsteady two-


NACA
RESEARCH ABSTRACTS NO. 66


dimensional supersonic flow round airfoils, that due
to Van Dyke is combined with the linearized super-
sonic theory for a rectangular flat plate to obtain the
effect on the damping in pitch. The values given
apply to wings of symmetrical biconvex section
oscillating at low frequency, in a stream of Mach
number such that the shock at the leading edge is
attached.




MISCELLANEOUS




Errata No. 1 on "THIRTY-SEVENTH ANNUAL
REPORT OF THE NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS REPORT 1028."
1951.





UNPUBLISHED PAPERS





N-31692*

ON THE CRITERION OF WING TORSIONAL STIFF-
NESS. (Sul Criterio Di Rigidezza Torsionale Delle
Strutture Alari). Placido Cicala. February 1954.
15p. diagrs. (Trans. from Onore di Modesto
Panetti, 1950, p. 45-49)

The official regulations for the stress analysis of
airplanes provide criteria which specify in simple
form the required torsional stiffness. At the time
the Italian standards were established a criterion of
stiffness had been introduced which consisted in
fixing the maximum angle of torsion produced by a
stipulated torsional moment, distributed spanwise by
a law proportional to the square of the chord. The
present report is an attempt to improve this criterion
with due consideration to the characteristics of every
type of aircraft in relation to the threats of the cited
aeroelastic phenomena, while maintaining the
necessary simplicity.


NACA-Langley 7-13-54 4M




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