Research abstracts

National Advisory Committee For Aeronautics


Research Abstracts


NO.40


MARCH 27, 1953


CURRENT NACA REPORTS

NACA TN 2867

HEAT AND MOMENTUM TRANSFER BE FN A
SPHERICAL PARTICLE AND AIR STORE .
Tang, J. M. Duncan and H. E. Schweyer, Uinik
of Florida. March 1953. 48p. diagrs., hotto., tab
(NACA TN 2867)\

Heat-transfer coefficients for a spherical patLcle
heated by an induction coil in a moving air stream
were experimentally determined for the Reynolds
number range from 50 to 1000 using spheres of 1 8-
to 5 8-inch diameter and air velocities from 1 to 13
feet per second. A correlation of the heat-transfer
factor or Stanton number with the Reynolds number
was obtained and expressed by an empirical equation.
This correlation is in agreement with the values
calculated from theory for the lower range of
Reynolds numbers studied. The skin-friction factor
representing the momentum transfer calculated from
the boundary-layer theory shows good agreement with
the experimental heat-transfer factor except in the
lower range of Reynolds numbers studied The re-
lationship St = Cf 2 where St is the Stanton num-
ber and Cf is the skin-friction factor is suggested
for the case of an air stream flowing around a sphere.
An empirical equation relating the heat-transfer
factor to the total-drag coefficient is also suggested.

NACA TN 2890

A LINEAR TIME-TEMPERATURE RELATION FOR
EXTRAPOLATION OF CREEP AND STRESS-
RUPTURE DATA. S S. Manson and A. M. Haferd.
March 1953. 49p. diagrs. (NACA TN 2890)

A time-temperature parameter based on examination
of the published stress-rupture data for a variety of
materials is proposed in the form
(T Tg) (log t log ta), where T is temperature in
degrees Fahrenheit, t is rupture time or the time
to obtain a given total creep elongation, and Ta and
log ta are material constants which appear to be
determinable from suitable rupture data in the time
range between 30 and 300 hours For the 40 mate-
rials investigated, use of this parameter in conjunc-
tion with experimental data involving rupture times
below 300 hours resulted in very good extrapolations
to longer rupture times (up to 10, 000 hr where data
were available) compared with corresponding predic-
tions obtainable from a recently proposed parameter
(T + 460)(20 + log t). For correlation of minimum
creep-rate data, the parameter used is in the form
(T Ta) (log r + log ra), where r is the creep rate
and ra is a material constant.


*AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST, NW
THE REPORT TITLE AND AUTHOR
4 'z._ (? 5
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NACA TN 2902


'/. MATRIX METHODS FOR DETERMINING THE
S., IONGITUDINAL-STABILITY DERIVATIVES OF AN
-APLANE FROM TRANSIENT FLIGHT DATA.
SJames J. Donegan. March 1953. 65p. diagrs., 6
5) tabs (NACA TN 2902)

.'Three methods are presented for calculating the lon-
/ gitudinal stability derivatives from transient flight
data. Several examples using flight data are given
to illustrate the method. The results indicate the
scatter which is typical of this type of analysis.

NACA TN 2903

IMPINGEMENT OF CLOUD DROPLETS ON AERO-
DYNAMIC BODIES AS AFFECTED BY COMPRESS-
IBILITY OF AIR FLOW AROUND THE BODY
Rinaldo J. Brun, John S. Serafini and Helen M.
Gallagher March 1953. 20p. diagrs. (NACA
TN 2903)

The trajectories of water droplets in a compressible-
air flow field around a cylinder were computed with a
mechanical analog. The results of the calculations at
the flight critical Mach number were compared with
calculations of trajectories in an incompressible
flow field. For a cylinder, the effect of compress-
ibility of the air on the droplet trajectories was
negligible up to the flight critical Mach number.
The results obtained with the cylinder were extended
to airfoils. This extension is possible because the
incompressible flow fields of both cylinders and
airfoils are similarly altered by compressibility.

NACA TN 2904

IMPINGEMENT OF WATER DROPLETS ON A CYL-
INDER IN AN INCOMPRESSIBLE FLOW FIELD AND
EVALUATION OF ROTATING MULTICYLINDER
METHOD FOR MEASUREMENT OF DROPLET-SIZE
DISTRIBUTION, VOLUME-MEDIAN DROPLET SIZE,
AND LIQUID-WATER CONTENT IN CLOUDS
Rinaldo J Brun and Harry W. Mergler. March
1953. 71p. diagrs photo., 4 tabs. INACA
TN 2904)

The trajectories of water droplets in an incompressi-
ble flow field around a cylinder were calculated with
a mechanical analog The collection efficiency, the
area of droplet impingement on the cylinder, and the
rate of droplet impingement were determined from
the trajectories An evaluation of the rotating multi-
cylinder method for the measurement of droplet-size
distribution, volume-median droplet size, and
liquid-water content was made based on the results
of the trajectory calculations.


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







2


NACA TN 2910

AN APPLICATION OF THE METHOD OF CHARAC-
TERISTICS TO TWO-DIMENSIONAL TRANSONIC
FLOWS WITH DETACHED SHOCK WAVES. Keith
C. Harder and E. B. Klunker. March 1953. 16p.
diagrs. (NACA TN 2910)

An application of the method of characteristics
presented which affords a means for determining the
surface pressures for a class of two-dimensionk
airfoils of given nose shape and arbitrary rear art
in a sonic or supersonic stream if surface press ure
data are given for one member of the class. Fo
engineering purposes, the method of characters ics
may be replaced by a simple application of Pran tl-
Meyer flow concepts. An explanation of the nonlinear
force characteristics of two-dimensional airfoils at
transonic speeds is presented on the basis of sensi-
tivity of these flows to changes in geometry and angle
of attack.

NACA TN 2911

A LOW-SPEED EXPERIMENTAL STUDY OF THE
DIRECTIONAL CHARACTERISTICS OF A SHARP-
NOSED FUSELAGE THROUGH A LARGE ANGLE-
OF-ATTACK RANGE AT ZERO ANGLE OF SIDE-
SLIP. William Letko. March 1953. 27p. diagrs.,
photo. (NACA TN 2911. Formerly RM L52J14)

Static yawing moments and some instantaneous
yawing moments are presented through a range of
angle of attack at zero sideslip angle for a plain
fuselage model having a sharp conical nose and for
the fuselage with several nose modifications, one
of which consisted of a ring located at different sta-
tions along the nose. Some circumferential pressure
measurements at one station on the body at several
angles of attack are also presented.


NACA TN 2912

THE NORMAL COMPONENT OF THE INDUCED
VELOCITY IN THE VICINITY OF A LIFTING ROTOR
AND SOME EXAMPLES OF ITS APPLICATION.
Walter Castles, Jr. and Jacob Henri De Leeuw,
Georgia Institute of Technology. March 1953. 38p.
diagrs., 3 tabs. (NACA TN 2912)

A method is presented for computing the approxi-
mate values of the normal component of the induced
velocity at points in the flow field of a lifting rotor.
Tables and graphs of the relative magnitudes of the
normal component of the induced velocity are given
for selected points in the longitudinal plane of
symmetry of the rotor and on the lateral rotor axis.
A method is also presented for using the tables and
graphs to determine the interference induced
velocities arising from the second rotor of a tandem-
or side-by-side-rotor helicopter and the induced flow
angle at a horizontal tail plane.


NACA TN 2913

ON THE DEVELOPMENT OF TURBULENT WAKES
FROM VORTEX STREETS. Anatol Roshko,


NACA
RESEARCH ABSTRACTS NO.40


California Institute of Technology. March 1953.
77p. diagrs., photos., 3 tabs (NACA TN 2913)

Wake development behind circular cylinders at \
Reynolds numbers'from 40 to 10,000 was investigated'
by hot-wire techniques in a low-speed wind tunnel.
The Reynolds number range of periodic vortex
shedding is divided into two distinct subranges. In
the stable range, |R = 40 to 150, regular vortex
streets are formed and no turbulent motion develops,
the vortices decaying by viscous diffusion. The
range R = 150 to 00 is a transition region to the
irregular range i which turbulent velocity fluctua-
tions accompany he periodic formation of vortices.
The diffusion is t rbulent and the wake becomes
fully turbulent in 40 to 50 diameters The turbulence
is initiated by la inar-turbulent transition in the
free layers whici spring from the separation points
on the cylinder. An annular vortex street was
observed in the rake of a ring.

NACA TN 2914

A METHOD FO RAPID DETERMINATION OF THE
ICING LIMIT O A BODY IN TERMS OF THE
STREAM COND IONS. Edmund E. Callaghan and
John S. Serafin. March 1953. 33p. diagrs.
(NACA TN 2914)

The effects of existing frictional heating were ana-
lyzed to determine the conditions under which ice
formations on aircraft surfaces can be prevented. A
method is presented for rapidly determining by
means of charts the combination of Mach number,
altitude, and stream temperature which will maintain
an ice-free surface in an icing c;oud. The method
can be applied to both subsonic and supersonic flow.
The charts presented are for Mach numbers up to
1. 8 and pressure altitudes from sea level to 45, 000
feet.

I
NACA TN 2915

EFFECT OF PROCESSING VARIABLES ON THE
TRANSITION TEMPERATURE, STRENGTH, AND
DUCTILITY OF HIGH-PURITY, SINTERED,
WROUGHT MOLYBDENUM METAL. Kenneth C.
Dike and Roger A. Long. March 1953. 26p. diagrs.,
photos., 3 tabs. (NACA TN 2915)

High-purity, sintered, wrought molybdenum metal
has a transition temperature range near room tem-
perature which varies according to the amount of
swaging reduction, increasing swaging reduction de-
creases the transition temperature range. Recrys-
tallized metal possesses a higher transition temper-
ature range than as-swaged metal, regardless of the
amount of prior working. Ultimate tensile strengths
of as-swaged, stress-relieved, or recrystallized
metal are affected by varying amounts of working
from 35 to 99 percent swaging reduction, but the
differential is not large. Ductility of as-swaged
metal at room temperature increases with increased
working; however, when stress-relieved, the duc-
tility does not vary with working. Recrystallized
metal has good ductility above the transition range







NACA
RESEARCH ABSTRACTS NO.40

provided that the prior working is greater than about
50 percent. Other data are presented as to grain
size, type of fracture, and chemical analysis.


NACA TN 2917

A MODIFIED REYNOLDS ANALOGY FOR THE COM-
PRESSIBLE TURBULENT BOUNDARY LAYER ON A
FLAT PLATE. Morris W. Rubesin. March 1953.
23p. diagrs., tab. (NACA TN 2917)

A modifed Reynolds analogy is developed for the
compressible turbulent boundary layer on a flat
plate. When mixing-length theories are used to
evaluate terms of the final expressions, it is found
for air that the ratio of Stanton number to half the
local skin-friction coefficient is greater than unity.
At Mach number equals zero, this ratio is of the
order of 1.18 to 1.21 for Reynolds numbers based on
momentum thickness of 103 to 106. Up to a Mach
number of 5 and under extreme conditions of surface
temperature, it is found that the ratio of Stanton num-
ber to half the skin-friction coefficient differs from
its values for the incompressible case (M = 0) by
amounts so small as to be of the magnitude of the
uncertainties in the theory.


NACA TN 2918

EFFECTS OF PARALLEL-JET MIXING ON DOWN-
STREAM MACH NUMBER AND STAGNATION PRES-
SURE WITH APPLICATION TO ENGINE TESTING IN
SUPERSONIC TUNNELS. Harry Bernstein. March
1953. 26p. diagrs., photos. (NACA TN 2918)

A one-dimensional analysis of the results of the
parallel-jet mixing encountered in the testing of en-
gines in supersonic wind tunnels is reported. This
type of analysis presents a reasonable approach to
obtairung approximate figures for the tunnel operating
conditions while the tunnel is still in the design stage.
These figures would be based upon the known tunnel
geometry.and inlet conditions and estimations of the
model geometry and values of the engine performance
parameters. Additional equations are presented for
evaluation of changes due to the burning of excess
fuel downstream of the engine-exhaust station.



NACA TM 1355

STUDY OF THE SUPERSONIC PROPELLER. (Etude
de L'Helice Supersonique). Jean Fabri and Raymond
Siestrunck. March 1953. 23p. diagrs. (NACA
TM 1355. Trans. from Ministere de 1'Air. Publica-
tions Scientifiques et Techniques 248, 1951, p. 113-
130; International Conference on Mechanics,
Proceedings, v. 1, 1950).

In this paper a propeller having all sections operating
at supersonic speeds is designated a supersonic pro-
peller regardless of flight speed. Analyses assume
subsonic flight speeds but very high rotational speeds.
A very elementary analysis of the efficiency of a jet-
propeller system is presented. A propeller analysis
based on conventional vortex blade element theory is


3


presented and reduced to a single point method which
leads to an expression for optimum advance ratio in
terms of hub-tip diameter ratio and airfoil fineness
ratio. An expression for propeller efficiency in
terms of advance ratio, hub-tip diameter ratio, and
airfoil thickness ratio is also presented. Use is
made of theoretical airfoil characteristics at super-
sonic speeds. A study of blade section interference,
blade shock and expansion fields, at supersonic sec-
tion speeds is presented. An example taken indicates
that an efficiency of seventy percent can be obtained
with a propeller having a tip Mach number of 2.3.


BRITISH REPORTS


N-21075*

Royal Aircraft Establishment (Gt. Brit.)
COMMUNICATION IN THE PRESENCE OF NOISE.
D. J. Richardson. October, 1952. 29p. (RAE
Tech. Note GW 215)

This report is divided into three main parts. The
first part considers discrete information, and a
number of theorems on discrete information are
developed. The second discusses continuous infor-
mation, and shows how the various theorems on
discrete information may be generalized to apply to
the continuous case. An information potential
measure is introduced. The third is concerned with
the transfer of continuous information in the pres-
ence of noise.



N-21206*

Aeronautical Research Council (Gt. Brit.)
THE CALCULATION OF AERODYNAMIC DERIVA-
TIVE COEFFICIENTS FOR WINGS OF ANY PLAN
FORM IN NON-UNIFORM MOTION. W. P. Jones.
1952. 12p. diagrs., tab. (ARC R& M 2470.
Formerly ARC 10, 142; 0.623; FM 1030; S & C 2074)

In the present paper, a method is outlined for the cal-
culation of aerodynamic forces on wings of any plan
form in steady or unsteady motion. It is based on
vortex sheet theory for thin wings as developed by
the writer in previous reports, but makes use of
Falkner's scheme for the approximate calculation of
downwash distributions. In the earlier work, satis-
factory agreement with the experimental evidence
available was obtained, but as the downwash distribu-
tions were calculated exactly the numerical work was
rather complicated and involved the use of incomplete
elliptic integrals and the treatment of singularities.
The method now proposed avoids these computational
difficulties and is perhaps more suitable for routine
calculations of flutter derivatives. Satisfactory
solutions of many problems in steady motion have
already been obtained by Falkner using approximate
downwash distributions as determined by his vortex
lattice, and it is thought that the similar scheme
suggested in this paper might also prove to be suffi-
ciently accurate for problems in unsteady motion.
The computational procedure is briefly summarized







4


in appendix I. Certain modifications of the main
scheme which would further reduce the amount of
computation are indicated in the main text.


N-21207 *

Aeronautical Research Council (Gt. Brit.)
PART I. TABULATED THERMAL DATA FOR
HYDROCARBON OXIDATION PRODUCTS AT HIGH
TEMPERATURES. PART II. THE EFFECT OF
DISSOCIATION ON ROCKET PERFORMANCE CAL-
CULATIONS. A. B. P. Beeton. 1952. 14p.
diagrs., 3 tabs. (ARC R & M 2542. Formerly
RAE Tech. Note Aero 1835; SD 40. RAE Tech.Note
Aero 1838; SD 42)

Tables are given of the total heat and entropy of H2O,
CO2, 02, CO, H2, OH, O, and H for the range of
temperature 15000-40000 K. Values are also given
for the corresponding equilibrium constants over the
same temperature range. The tables have been
compiled with a view to their use in calculating the
performance of liquid-fuel rockets.


N-21208 *

Aeronautical Research Council (Gt. Brit.)
FLIGHT TESTS ON HURRICANE II, Z. 3687 FITTED
WITH SPECIAL WINGS OF 'LOW-DRAG' DESIGN.
R. H. Plascott, D. J. Higton, F. Smith and A. R.
Bramwell. 1952. 13p. diagrs. (ARC R& M 2546;
ARC 9172; ARC 10, 106. Formerly RAE Aero 2090;
Aero 2153)

This report describes flight tests to investigate the
profile-drag characteristics of a "low-drag" section
wing built by Armstrong Whitworth, Ltd., using a
new type of construction of their own design. During
the first series of tests, a section of the wing was
pressure-plotted and the results showed that it should
be possible to obtain laminar flow over a range of
lift coefficient from 0. 12 to 0. 50. A few preliminary
profile-drag measurements were also made and a
fairly low profile-drag coefficient (CD = 0. 0046 to
0. 0050) was recorded over a lift coefficient range of
0. 20 to 0. 40; there was, however, a rapid rise in the
profile drag coefficient at lift coefficients less than
0. 20, and investigation of the surface waviness
showed that the failure to maintain laminar flow at
higher speeds was probably due to the excessive
waviness present, which amounted to a variation of
about 2-1/2 thousandths of an inch from the mean
deflection curve on a two-inch gage length. A
further series of profile-drag measurements was
made when the surface waviness had been reduced
to 1 thousandth of an inch variation from the mean
deflection curve on a 2-inch gage length. It was
found that, provided no flies or other insects were
picked up during the flight, the drag coefficient had
been reduced to 0. 0044 over a range of lift coeffi-
cient from 0. 12 to 0.50. This corresponds to tran-
sition from 50 to 60 percent chord. With the
reduced surface waviness, it was possible to main-
tain laminar flow up to Reynolds numbers of nearly
20 millions.


NACA
RESEARCH ABSTRACTS NO.40


N-21209*

Aeronautical Research Council (Gt. Brit.)
TESTS IN THE COMPRESSED AIR TUNNEL ON THE
AEROFOILS NACA 0015 AND NACA 0030 WITH AND
WITHOUT SPLIT FLAP AND ON OTHER AEROFOILS
OF VARIOUS THICKNESSES WITH A SPLIT FLAP.
R. Jones. 1952. 32p. diagrs., 16 tabs. (ARC
R & M 2584. Formerly ARC 4191; ARC 4511;
ARC 4607)

This report contains the results of experiments to
determine the effect of thickness on the aerodynamic
characteristics of a group of rectangular wings with
and without a split flap. Wings with the following
airfoil sections were tested: NACA 0015 and NACA
0030 with and without a split flap; NACA 0012, NACA
23012, RAF 28 and RAF 48 with the flap. The effect
of rounding the edge of the flap was considered on the
wing with NACA 0015 section. The effect of round-
ing the ends of the NACA 0030 section was examined.
CL, CD, and Cm were obtained over a range of
Reynolds numbers with additional CD measurements
at closer intervals of R on the two wings without flap.


N-21210*

Aeronautical Research Council (Gt. Brit.)
THE EFFICIENCY OF A PITOT INTAKE INCLINED
TO THE AIR STREAM. E. L. Place and R.
Lecavalier. 1952. 8p. diagrs. (ARC R& M 2621;
ARC 11, 162. Formerly NGTE Memo. M. 21)

In an earlier report on intake ducting for supersonic
flight, the efficiency of a "pitot" type intake was dis-
cussed and shown to have a marked effect on the
performance of gas-turbine engines. The present
report is supplementary in that it describes the ef-
fect of inclining the pitot intake to the main air-
stream direction in the transonic Mach number range
0. 7 to 1. 5, an effect which is at present incalculable.
Curves are presented showing the influence of incli-
nation on intake adiabatic efficiency and air mass
flow into the intake. These experimental results are
then illustrated by application to the performance of
a typical turbine engine and a propulsive duct in
sonic and supersonic flight. At a flight Mach number
of 1. 5, it is found that, for both turbine engine and
propulsive duct, an inclination of 50 reduces the net
thrust by roughly 2 percent compared with the nor-
mal flight thrust. For inclinations greater than 50,
however, thrust falls off more rapidly, and at 100
inclination, it is reduced by roughly 6. 5 percent for
the turbine engine and 7. 5 for the propulsive duct.


N-21211*

Aeronautical Research Council (Gt. Brit.)
REGENERATOR HEAT EXCHANGERS FOR GAS-
TURBINES. J. E. Johnson. 1952. 72p. diagrs.,
13 tabs. (ARC R & M 2630; ARC 11, 770. Formerly
RAE Aero 2266; SD 27)

Information was required from which the perform-
ance of regenerators suitable for heat exchangers for
gas turbines could readily be estimated. A series of







NACA
RESEARCH ABSTRACTS NO.40


tables and curves have been prepared from which the
efficiency of a regenerator can be calculated if the
operating conditions and heat-transfer coefficients
are known. The tables and curves cover a range of
lengths and blow times appropriate to gas-turbine
conditions. Measurements of heat transfer and
pressure drop coefficients have been made on sev-
eral examples of matrix of both the gauze and flame
trap type in conditions similar to those in a gas
turbine. A number of examples have been worked
out from the experimental results to show the rela-
tive importance of the different variables on the
performance of typical regenerators. A gauze
matrix of fine wire and open mesh has a much lower
weight and only slightly higher pressure drop than a
flame-trap matrix for the same efficiency. The
recommended size of gauze is a wire diameter of
0.002 in. to 0.004 in. and a mesh of 20 to 40 wires
per inch; the material should be stainless steel.
Further design study is necessary to determine
whether this advantage can be maintained in a com-
plete regenerator.


N-21212*

Aeronautical Research Council (Gt. Brit.)
COMPARATIVE FLUTTER TESTS ON TWO, THREE,
FOUR AND FIVE-BLADE PROPELLERS. H. G.
Ewing, J. Kettlewell and D. R. Gaukroger. 1952.
8p. diagrs. (ARC R& M 2634; ARC 11,438.
Formerly RAE Structures 18)

This report describes comparative flutter tests on
two-, three-, four-, and five-blade Duralumin
propellers with the same blade design. The tests
were made on the No. 3 spinning tower, Royal Air-
craft Establishment. Strain gages were used for
determining the vibratory stresses and the phase
relations between the blades. A wide range of blade
angles above and below the stalling region was
explored. Stalling flutter was the only form encoun-
tered. The phase relation of the blades was found to
be dependent on number of blades and speed of rota-
tion, and to influence the amplitude of the vibratory
stresses. It is shown that no direct comparison of
the flutter characteristics of the two-, three-, four-,
and five-blade propellers can be made.


N-21213*

Aeronautical Research Council (Gt. Brit.)
CONCERNING THE ANNULAR AIR INTAKE IN
SUPERSONIC FLIGHT. I. M. Davidson and L. E.
Umney. 1952. 22p. diagrs., photos. (ARC
R & M 2651; ARC 11, 645. Formerly NGTE R. 16)

The stability of an annular air intake at a Mach num-
ber of 1.4 and with Reynolds numbers of about
1. 5 x 106 is considered in detail and a method is de-
scribed whereby the experimental results might be
extrapolated for preliminary full-scale design pur-
poses. This extrapolation has yet to be checked
experimentally, but suggests that a typical aircraft
intake would have an over-all isentropic efficiency of
about 85 percent. The results also indicate that both
the stability and the efficiency of an intake could be


5


improved by controlling the boundary layer on its
nacelle, and as an alternative to boundary-layer
suction a device which is described as a segregation
ring is suggested. This, it appears, might raise the
efficiency by some 2 or 3 percent.

N-21214*

Aeronautical Research Council (Gt. Brit.)
THE PHYSICAL CHARACTERISTICS OF WIRE
RESISTANCE STRAIN GAUGES. Eric Jones and
K. R. Maslen. 1952. 44p. diagrs., photos., 4
tabs. (ARC R & M 2661; ARC 12,357. Formerly
RAE Instn. 2)

This report deals with the fundamental principles of
the wire resistance strain gage. Types of strain
gage in common use and their methods of construc-
tion are described, and the mechanism whereby
strain effects change of resistance is discussed. A
subsection is devoted to the behavior of fine wires,
in general, under strain. Possible causes of error,
including the effects of humidity and temperature,
are discussed, and as far as possible methods are
given of overcoming these difficulties. The effect
of the passage of current on the strain gages is
described, and methods of increasing the output are
suggested. The final section is devoted to miscel-
laneous properties of the wire resistance strain
gage, on several of which very little information is
at present available.

N-21215*

Aeronautical Research Council (Gt. Brit.)
THEORETICAL INVESTIGATIONS OF TERNARY
LIFTING SURFACE CONTROL SURFACE -
TRIMMING TAB FLUTTER AND DERIVATION OF A
FLUTTER CRITERION. H. Wittmeyer. 1952.
42p. diagrs., 3 tabs. (ARC R &M 2671; ARC
12,043. Formerly RAE Structures 19)

Theoretical investigations have been made of the
flutter of an idealized trimming tab system having
three degrees of freedom normal translation of the
main lifting surface, rotation of the control surface,
and rotation of the tab. All the structural param-
eters of the system have been varied except the out-
of-balance moment of the control surface. The
cases in which the system is free from flutter have
been particularly investigated. From these investi-
gations, criteria for the avoidance of flutter have
been derived. If the structural parameters of
the system satisfy these criteria, flutter of the
system with these three degrees of freedom should
be impossible. The results are applicable to trim-
ming tabs, servotabs with zero follow-up ratio, and
generally to all systems in which the tab can be
regarded as connected elastically only to the control
surface.

N-21217'

Aeronautical Research Council (Gt. Brit.)
FLOW THROUGH A HELICOPTER ROTOR IN VER-
TICAL DESCENT. P. Brotherhood. 1952. 14p.
diagrs., photos. (ARC R & M 2735; ARC 11,837.
Formerly RAE Aero 2272)








6


Flight tests have been made on a Hoverfly I helicop-
ter to investigate the types of flow associated with
various rates of vertical descent. At the same time,
measurements of the performance were made. The
results are analyzed by two different methods to pro-
duce characteristic curves for the rotor and are
compared with data obtained from wind-tunnel tests
on model propellers at negative rates of advance.
The information was obtained from the Hoverfly I
helicopter but it is thought that the results can be
applied to any other helicopter of similar size.

N-21218*

Aeronautical Research Council (Gt. Brit.)
LOW-SPEED WIND-TUNNEL TESTS OF FOWLER
FLAPS, SLATS AND NOSE FLAPS ON A MODEL OF
A JET AIRCRAFT WITH A 40 DEG SWEPT-BACK
WING. A. Spence. 1952. 18p. diagrs., 6 tabs.
(ARC R & M 2752; ARC 12,131. Formerly RAE
Aero 2302)

This report presents the results of tests with Fowler
flaps on a model of a single-jet aircraft with a 400
sweptback 10-percent-thick wing. Slats and nose
flaps were also tested as means of delaying the tip
stall. The maximum trimmed lift coefficient without
flaps or slats was 1.055 (R = 2.7 x 106). With half-
span Fowler flaps (leaving a gap across the fuselage)
and slats over the outer half of the span, this value
was increased to 1.64, and there was adequate
stability. Tests in which the spanwise extent of the
nose flap was varied, indicated that about 50 percent
wing semispan per side was the optimum length of
slat or nose flap for avoiding instability at the stall.

N-21219*

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE INFLUENCE OF ASPECT RATIO ON
THE VARIATION WITH MACH NUMBER OF THE
LIFT AND HINGE-MOMENT CHARACTERISTICS OF
A WING AND FULL-SPAN CONTROL. A. D. Young
and P. R. Owen. 1952. 6p. diagr., tab. (ARC
R & M 2767; ARC 7046; ARC 7133. Formerly RAE
Tech. Note Aero 1250; RAE Tech. Note Aero 1263)

It is shown on the basis of the linearized theory that
the effects of compressibility on the lift and hinge-
moment characteristics of a wing and full-span con-
trol are functions of aspect ratio. With reduction in
aspect ratio, the increase of the lift characteristics
with Mach number is reduced appreciably. The
same effect is noted for the hinge-moment character-
istic bl. The effects on the hinge-moment char-
acteristics b2 and b3 are rather more complicated,
but in many practical cases the influence of aspect
ratio will be very small.


N-21258*

Royal Aircraft Establishment (Gt. Brit.)
THE YOUNG'S MODULUS, POISSON'S RATIO AND
RIGIDITY MODULUS OF SOME ALUMINIUM
ALLOYS. PART I. N. Dudzinski. November
1952. 28p. photos., 14 tabs. (RAE Met. 69)


NACA
RESEARCH ABSTRACTS NO.40


Various binary and ternary chill-cast or sintered
aluminium alloys were prepared and their elastic
properties investigated. It was found that the
nitrides of aluminium, chromium, magnesium, iron,
vanadium, titanium, and zirconium caused an appre-
ciable increase in Young's modulus. Ternary inter-
metallic compounds showed a similar effect. The
specific Young's modulus value was improved by the
additions of chromium or manganese. The rigidity
modulus value in torsion was similar to that calcu-
lated from E and Poisson's ratio. The effect of
composition and the heat of formation of various
.intermetallic compounds is discussed.


N-21259 *

Ministry of Supply (Gt. Brit.)
ALUMINIUM-COPPER-CADMIUM SHEET ALLOYS.
H. K. Hardy. May 22, 1952. 25p. diagrs.,
photos., 13 tabs. (MOS S & TM 9; 52; Fulmer
Research Institute Ltd.)

Aluminum-copper-cadmium alloys have been rolled
to sheet on a semitechnical scale and showed an
exceptional capacity for hot work. The aging char-
acteristics have been examined and an uncoated alloy
with 5 percent copper-0. 15 percent cadmium gave
25 tons/sq in. for the 0. 1 percent proof stress and
30 tons/sq in. maximum stress in both the trans-
verse and longitudinal directions. Cold working
prior to artificial aging markedly raised the tensile
properties of aluminum-copper alloys but had the
unique effect of reducing the properties of the
aluminum-copper-cadmium alloys. The proof stress
can drop by 3 tons/sq in. but the effect can be
prevented if the alloy is given a short partial aging
("incubation") treatment at 1650 or 1700 C before
cold working. Aluminum-copper alloys showed a
marked susceptibility to stress corrosion resulting
from severe intercrystalline corrosion. Aluminum-
copper-cadmium alloys were immune from stress
corrosion and showed an intracrystalline attack when
sprayed with 3 percent sodium chloride solution.


N-21262 *

Aeroplane and Armament Experimental Establish-
ment (Gt. Brit.) AN INVESTIGATION INTO THE
PITOT RAKE METHOD OF MEASURING TURBO JET
ENGINE THRUST IN FLIGHT. J. Stephenson, R. T.
Shields and D. W. Bottle. December 23, 1952.
56p. diagrs., 9 tabs. (AAEE/Res 265)

Tests have been made to establish whether a pitot
rake could be used as an absolute measure of the
thrust of a jet engine on the ground and in flight.
The tests were made to investigate errors due to the
assumptions inherent in the single pilot method of
estimating thrust in flight, and to establish if a rake
can be used to calibrate the single pitot of an uncal -
ibrated engine installed in an aircraft. The tests
were also planned to check the generally accepted
nondimensional thrust relationship for jet engines.
The tests were made on Derwent 5 engines installed
in Meteor 4 aircraft. The tests covered a wide
range of flight conditions and included test bed







NACA
RESEARCH ABSTRACTS NO. 40

measurements on bare engines and later, ,measure-
ments of exit static pressure. Although the tests
could not all be made on the same engine, the same
final nozzle was used in all the main tests. The
main conclusions were: (1) Static tubes must be
Incorporated in the pitot rake to give absolute thrust
measurements and even so a discrepancy of 2 percent
requires further investigation. (2) The single pitot
method of estimating flight thrust based or engine
test bed calibration was in error by as much as 6
percent due to changes in total pressure sampled by
the single pitot and in the magnitude of the exit static
pressure between calibration and test conditions.
Using the pitot static rake to calibrate the single pitot
of an installed engine introduced flight thrust errors
no larger and possible smaller. (3) Nondimensional
thrusts at 35, 000 ft were some 7 percent lower than
corresponding thrusts at 5, 000 ft. Further tests are
required to establish the magnitude of these effects
on other engine types.

N-21311"

Aeronautical Research Council (Gt. Brit.)
EFFECT OF THE CONTROL-CIRCUIT ON
FLUTTER. (Einfluss der rudersteiirung auf das
flattern). K. Leiss. June 17, 1948. 50p. diagrs.
(ARC 11,583; 0.734. Trans. from Zentrale fir
wissenschaftliches Berichtswesen der
Luftfahrtforschung, Berlin. FB 1670, October 10,
1942)

As a result of the participation of the controls, the
following types of flutter appear: when the control
is slack there is flutter of the system with the con-
trol organ and the control surface out of phase, but
when the control is stiff there is flutter with the two
components in phase. The second type of flutter, in
particular, has only been treated incompletely, if at
all, up to the present; but because of its appearance
the advantage of stiffening the control is doubtful.
Mass-balancing of the control surface, and with
occasional exceptions structural damping in the
control system, improve the flutter properties,
especially when the control organ and the control
surface are in phase with each other. The oscil-
lation of the controls is included in the flutter cal-
culations by means of suitable transformations;
this is done in such a way that little more computa-
tional work is involved than for the case where the
oscillation of the control system is neglected.


N-21394 *

National Gas Turbine Establishment (Gt. Brit.)
THE BURNING OF SINGLE DROPS OF FUEL.
PART m. COMPARISON OF EXPERIMENTAL AND
THEORETICAL BURNING RATES AND DISCUSSION
OF THE MECHANISM OF THE COMBUSTION PROC-
ESS. G. A. E. Godsave. August 1952. 47p.
diagrs., photos. (NGTE R. 88)

A theoretical and experimental investigation has been
made of the factors influencing the rate of decrease
in size of a single burning drop, as in a burning fuel
spray. The experimental technique adopted in this
investigation has been to suspend a single drop of


7

fuel on a fine silica filament, and to measure with a
cinematograph camera the rate of decrease in size
during combustion in still air. A theoretical treat-
ment given in part I indicates that the heat transfer
to the drop is the dominant factor determining its
life. The coefficient of heat transfer to the drop is
itself a function of the mass evaporation rate. The
analysis given enables a solution to be obtained for
the heat transfer to the drop in the presence of the
evaporative flow of the vapor. The experimental
results presented in part II have been examined here
in part III on the basis of such a heat-transfer rela-
tionship, with due allowance for any significant
radiation effects. A discussion is given of the
mechanism of the combustion process during the
burning of a fuel spray.


MISCELLANEOUS


NACA TN 2278

Errata No. 1 on "THEORETICAL SYMMETRIC SPAN
LOADING DUE TO FLAP DEFLECTION FOR WINGS
OF ARBITRARY PLAN FORM AT SUBSONIC
SPEEDS." John DeYoung. January 1951.


NACA TN 2770

Errata No. 1 on "STUDY OF THE PRESSURE RISE
ACROSS SHOCK WAVES REQUIRED TO SEPARATE
LAMINAR AND TURBULENT BOUNDARY LAYERS."
Coleman duP. Donaldson and Roy H. Lange.
September 1952.


UNPUBLISHED PAPERS


N-19881*

DIRECT-FORCE MEASUREMENTS OF TURBULENT
SKIN FRICTION ON CYLINDERS IN AXIAL FLOW
AT SUBSONIC AND SUPERSONIC VELOCITIES.
Dean R. Chapman and Robert H. Kester. 1953.
20p. diagrs., photos., tab. (To be presented at
annual Institute of the Aeronautical Sciences meeting,
New York, January 26-29, 1953.)

The principal results of a study of all known theories
for calculating turbulent skin friction in compressi-
ble flow are briefly reviewed. These various
theories predict widely different effects of Mach
number on skin friction, and hence also on heat
transfer. Systematic experiments were made to
determine the magnitude of turbulent skin friction
along the cylindrical portion of cone-cylinder bodies
of revolution having over-all fineness ratios of 10,
15, and 25. Data were obtained by directly measur-
ing forces. Boundary-layer surveys were made to
determine the correction necessary to apply to the
force measurements in order to determine the ef-
fective starting position of the turbulent flow. Mach
numbers between 0. 5 and 3.6, and Reynolds numbers
between 4 million and 32 million were investigated.
At a Mach number of 2.0, data were obtained (by




UNIVERSITY OF FLORIDA

II II 11 l lll PNACA
3 1262 08153 280 5 RESEARCH ABSTRACTS NO.40

distorting the flexible-plate walls of the wind tunnel)
for three different pressure distributions in order to
evaluate the effect of a moderate pressure gradient
on turbulent skin friction.


DECLASSIFIED NACA REPORTS


NACA RM E52F06

MECHANISM OF START AND DEVELOPMENT OF
AIRCRAFT CRASH FIRES. I. Irving Pinkel, G.
Merritt Preston and Gerard J. Pesman. August 28,
1952. ii, 97p. diagrs., photos., 2 tabs. (NACA
RM E52F06) (Declassified from Restricted, 2/27/53)

Full-scale aircraft crashes were made to investigate
the mechanism of the start and development of air-
craft crash fires. The results are discussed herein.
This investigation revealed the characteristics of the
ignition sources, the manner in which the combusti-
bles spread, the mechanism of the union of the com-
bustibles and ignition sources, and the pertinent
factors governing the development of a crash fire as
observed in this program.


NACA-Langley 3-27-53 4000

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