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National Advisory Committee for Aeronautics
DECEMBER 3, 1952
CURRENT NACA REPORTS
NACA Rept. 1050
FORMULAS FOR THE SUPERSONIC LOADING, LIFT
AND DRAG OF FLAT SWEPT-BACK WINGS WITH
LEADING EDGES BEHIND THE MACH LINES. Doris
Cohen. 1951. iii, 40p. diagrs. (NACA Rept. 1050)
The method of superposition of linearized conical
flows has been applied to the calculation of the aero-
dynamic properties, in supersonic flight, of thin flat,
swept-back wings at an angle of attack. The wings
are assumed to have rectilinear plan forms, with
tips parallel to the stream, and to taper in the con-
ventional sense. The investigation covers the
moderately supersonic speed range where the Mach --
lines from the leading-edge apex lie ahead of the wi. c
The trailing edge may lie ahead of or behind theK nh',
lines from its apex. The case in which the Nias "
from one tip intersects the other tip is not tre cS e
Formulas are obtained for the load dlstributial.0
total lift, and the drag due to lift. For the caeii ..
which the trailing edge is outside the Mach conefrom "-.
its apex (supersonic trailing edge), the formulas-are
complete. For the wing with both leading and tral)ng
edges behind their respective Mach lines, a degreeof ..
approximation is necessary. It has been found poas -
ble to give practical formulas which permit the total'" ..
lift and drag to be calculated to within 2 or 3 percent
of the accurate linearized-theory value. The local
lift can be determined accurately over most of the
wing, but the trailing-edge-tip region is treated only
approximately. Charts of some of the functions
derived are included to facilitate computing, and
several examples are worked out in outline.
NACA Rept. 1063
AIRFOIL PROFILES FOR MINIMUM PRESSURE
DRAG AT SUPERSONIC VELOCITIES-GENERAL
ANALYSIS WITH APPLICATION TO LINEARIZED
SUPERSONIC FLOW. Dean R. Chapman. 1952. ii,
14p. diagrs. (NACA Rept. 1063. Formerly TN 2264)
A derivation is presented of the basic equations which
determine the supersonic airfoil profile having mini-
mum pressure drag for certain prescribed structural
requirements. The basic equations are applicable to
a variety of practical structural requirements, and
can be used with either linear, second order, or
shock-expansion airfoil theory. A solution of the
basic equations is found in closed form using linear
airfoil theory. The results show that in most cases
the optimum profile has a blunt trailing edge. The
optimum distribution of thickness depends on the
Mach number, airfoil-thickness ratio, and base pres-
sure coefficient. The pressure drag of the optimum
profile is compared to that of the biconvex and
double-wedge profiles. A graphical method of
determining an optimum airfoil is developed and
applied to several examples.
NACA Rept. 1067
GENERALIZATION OF BOUNDARY-LAYER
MOMENTUM-INTEGRAL EQUATIONS TO THREE-
DIMENSIONAL FLOWS INCLUDING THOSE OF RO-
TATING SYSTEM. Artur Mager. 1952. ii, 16p.
diagrs. (NACA Rept. 1067. Formerly TN 2310).
Boundary-layer equations for application in three-
dimensional flows are developed. With the use of a
fixed velocity profile and an empirical friction law an
approximate solution in closed integral form is ob-
ed for a generalized boundary-layer momentum-
F t sickness and flow deflection at the wall in the
.l lIi case A numerical evaluation of this so-
lut- .7 i. rired out for data obtained in a curving non-
rouati~g duct shows a fair quantitative agreement
..,vith ire measured values.
NACA TN '2809
EXPERPIENTAL INVESTIGATION OF ECCEN-
T RICITY RATIO. FRICTION, AND OIL FLOW OF
""SHpQ T JOURNAL BEARINGS. G. B. DuBois and
--V. W. Ocvirk, Cornell University. November 1952.
79p. diagrs., photos., 4 tabs. (NACA TN 2809)
An experimental investigation was conducted to obtain
performance data on bearings of length-diameter
ratios of 1, 1/2, and 1/4 for comparison with theo-
retical curves. A 1. 375-inch-diameter bearing was
tested at speeds up to 6000 rpm and with unit loads
from 0 to 900 pounds per square inch. Experimental
data for eccentricity ratio and friction followed single
lines when plotted against a theoretically derived
capacity number, which is equal to Sommerfeld
number times the square of the length-diameter ratio.
The form of the capacity number indicates that under
certain conditions the eccentricity ratio is theoreti-
cally independent of bearing diameter. A method of
plotting oil flow data as a single line is shown.
Methods are also discussed for approximating a
maximum bearing temperature and evaluating the
effect of deflection or misalinement on the eccen-
tricity ratio at the ends of the bearings.
NACA TN 2813
THEORY AND PROCEDURE FOR DETERMINING
LOADS AND MOTIONS IN CHINE-IMMERSED HY-
DRODYNAMIC IMPACTS OF PRISMATIC BODIES.
Emanuel Schnitzer. November 1952. 51p. diagrs.
(NACA TN 2813)
*AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW.,
THE REPORT TITLE AND AUTHOR.
WASHINGTON 25, D. C., CITING CODE NUMBER ABOVE EACH TITLE;
A theoretical method is derived for computing the
motions and hydrodynamic loads during water land-
ings of prismatic bodies involving appreciable im-
mersion of the chines. A simplified method of com-
putation covering flat-plate and V-bottom bodies with
beam-loading coefficients greater than unity is given
as a separate section. Comparisons of theory with
experiment are presented as plots of impact lift
coefficient and maximum draft-beam ratio against
flight-path angle and as time histories of loads and
motions. Fair agreement is found to exist for chine-
immersed landings for angles of dead rise of 0 and
30, beam-loading coefficients from 1 to 36.5, flight-
path angles from 2 to 90, and trims from 6 to 45.
NACA TN 2815
A THEORETICAL INVESTIGATION OF THE EF-
FECT OF PARTIAL WING LIFT ON HYDHOODYNAM-
IC LANDINt CHARACTERISTICS OF V-BOTTOM
SEAPLANES IN STEP IMPACTS. Joseph L. Sims
and Emanuel Schnitzer. November 1952. 20p.
diagrs. (NACA TN 2815)
A theoretical investigation is made of the loads and
motions in water-landing impacts of wide prismatic
V-bottom seaplanes for constant partial wing-lift
conditions where the resultant velocity of the sea-
plane is normal to the keel. An approximate method
is given for applying the results of this in'.'esliIgatiin
to the more gent-r.lt] case of oblique impact. The
increase in vertical hdJr.,,J',,njnic load factor due
to wing-lift reduction is shown to be approximately
133 percent of the decrease in air load.
NACA TN 2816
WATER-PRESSURE DISTRIBUTIONS DUL RINu
LANDINGS OF A PRISMATIC MODEL HAVING AN
ANGLE OF DEAD RISE OF 22-1 /2 AND BEAM-
LOADN'. COEF FICIENTS OF 0.48 AND 0.97.
Robert F. Smiley. November 1952. 37p. diagrs.
6 tabs. (NACA TN 2816)
As part of an over-all program, smooth-water land-
ing tests of a prismatic float having an angle of dead
rise of 22-1/20 were made. Water-pressure, ve-
locity, draft, and acceleration data are presented.
Landings were made for beam-loading coefficients
of 0. 48 and 0. 97 at fixed trims between 0. 2 and
30. 3 for a range of flight-path angles from 4. 6 to
25. 9 and also for 90. The experimental pressure
distributions are found to be in fair agreement with
the predictions of the available theory; however,
better agreement is obtained by modification of the
NACA TN 2817
A THEORETICAL AND EXPERIMENTAL INVESTI-
GATION OF THE EFFECTS OF YAW ON PRES-
SURES, FORCES, AND MOMENTS DURING SEA-
PLANE LANDINGS AND PLANING. Robert F.
Smiley. November 1952. 98p. diagrs., 7 tabs.
(NACA TN 2817)
A theory for the side force, rolling moment, yawing
moment, and pressure distribution during yawed
landings and planing of seaplanes was developed. For
RESEARCH ABSTRACTS NO.33
the special case of the straight-sided wedge without
chine immersion, the results of the theoretical analy-
sis are presented in the form of generalized curves
covering all step landing conditions. Experimental
impact and planing data are presented for a prismatic
wedge having an angle of dead rise of 22.5 and are
shown to be in reasonable agreement with the theoret-
NACA TN 2819
EFFECT OF HIGH-LIFT DEVICES ON THE STATIC-
LATERAL-STABILITY DERIVATIVES OF A 45
SWEPTBACK WING OF ASPECT RATIO 4.0 AND
TAPER RATIO 0.6 IN COMBINATION WITH A BODY.
Jacob H. Lichtenstein and James L. Williams.
November 1952. 50p. diagrs., photos., 5 tabs.
(NACA TN 2819)
This paper contains the results of wind-tunnel tests
to determine the effects of plain and split trailing-
edge flaps, with and without leading-edge slats, on
the low-speed static-lateral-stability derivatives of
a 45 sweptback-wing-body configuration. Com-
parisons between the increments obtained from ex-
periment and those evaluated by simple sweep theory
in combination with measured lift and drag incre-
ments are also made.
NACA TN 2820
AN ANALYSIS OF THE ERRORS IN CURVE-FITTING
PROBLEMS WITH AN APPLICATION TO THE CAL-
CULATION OF STABILITY PARAMETERS FROM
FLIGHT DATA. Marvin Shinbrot. November 1952.
29p. diagrs., 2 tabs. (NACA TN 2820)
The problem of assessing the errors in the param-
eters obtained from a curve-fitting process is con-
sidered, and a scheme which may be applied toward
the solution of such problems is obtained. This
method is then specialized to the problem of finding
the errors in the calculated stability parameters of
an airplane, and an example is given.
NACA TN 2821
TORSION TESTS OF ALUMINUM-ALLOY STIFF-
ENED CIRCULAR CYLINDERS. J. W. Clark and
R. L. Moore, Aluminum Company of America.
November 1952. 38p. diagrs., photos., 2 tabs.
(NACA TN 2821)
Results are presented for the second series of tor-
sion tests on aluminum-alloy stiffened circular cyl-
inders, the first series having been reported in
NACA ARR 4E31. The cylinders were similar in
construction except that the wall thickness was
0. 020 inch for the first series and 0. 032 inch for
the second series. The significant observation
from both series of tests are summarized and some
comparisons are made with more recent theoretical
RESEARCH ABSTRACTS NO.33
NACA TN 2823
LANGLEY FULL-SCALE-TUNNEL INVESTIGATION
OF THE MAXIMUM-LIFT AND STALLING CHARAC-
TERISTICS OF A TRAPEZOIDAL WING OF ASPECT
RATIO 4 WITH CIRCULAR-ARC AIRFOIL SECTIONS.
Roy H. Lange. November 1952. 24p. diagrs.,
photos. (NACA TN 2823. Formerly RM L7H19)
Results are given of an investigation at high Reynolds
numbers and low Mach numbers to determine the
maximum-lift and stalling characteristics of a trape-
zoidal wing of aspect ratio 4 with 10-percent-thick,
circular-arc airfoil sections. The tests included
measurements of the lift, the drag, and the pitching-
moment coefficients of the basic wing and of the wing
with 0. 20-chord droop-nose and trailing-edge flaps
deflected both alone and in various combinations.
Scale effects were investigated at Reynolds number
ranging from 3.27 x 106 to 7.67 x 106.
NACA TN 2824
EFFECTS OF INDEPENDENT VARIATIONS OF
MACH NUMBER AND REYNOLDS NUMBER ON
THE MAXIMUM LIFT COEFFICIENTS OF FOUR
NACA 6-SERIES AIRFOIL SECTIONS. Stanley F.
Racisz. November 1952. 32p. diagrs., 2 tabs.
(NACA TN 2824)
The NACA 65-006, 64-009, 64-210, and 642-215
airfoil sections were tested in the Langley low-
turbulence pressure tunnel to determine the effects
of independent variations of Mach number and
Reynolds number on the maximum-lift character-
istics for the smooth and rough conditions. The
lift characteristics were determined at constant
Reynolds numbers ranging from 1. 5 x 106 to
9. 0 x 106 for Mach numbers which varied from 0. 1
to approximately 0. 5 for each Reynolds number.
NACA TN 2825
A COMPARATIVE EXAMINATION OF SOME MEAS-
UREMENTS OF AIRFOIL SECTION LIFT AND DRAG
AT SUPERCRITICAL SPEEDS. Gerald E. Nitzberg
and Stewart M. Crandall. November 1952. 30p.
diagrs. (NACA TN 2825)
Systematic trends in the lift- and drag-coefficient
variation with Mach number for a number of rela-
tively thick airfoil sections at moderately super-
critical speeds are pointed out. Shortcomings of
transonic similarity rules are discussed and a semi-
empirical correlation of drag data is presented.
Differences between the drag curves for various air-
foils are reduced by basing drag coefficient on total
pressure rather than dynamic pressure. The initial
supercritical loss of lift of airfoil sections having
supersonic flow over the upper surface only is shown
to result primarily from pressure changes on the
NACA TN 2826
SIMULATION OF LINEARIZED DYNAMICS OF GAS-
TURBINE ENGINES. J. R. Ketchum and R. T.
Craig. November 1952. 25p. diagrs., photo.
(NACA TN 2826)
A general method of simulation is presented, where-
by linearized dynamics of gas-turbine engines may
be simulated in most orderly fashion with greatest
economy of equipment and with most direct use of
engineering data. Correlation of experimental and
simulated responses of a turbojet engine are shown.
Use of generalization factors in determining the
coefficients necessary for simulation of engine
dynamics is discussed.
NACA TN 2827
INVESTIGATION OF A DIFFRACTION-GRATINu
INTERFEROMETER FOR USE IN AERODYNAMIC
RESEARCH. James R. Sterrett and John R. Erwin.
November 1952. 36p. photos., diagrs. (NACA
A low-cost interferometer that is easy to adjust and
has a large field of view is described. This instru-
ment, which is based on a principle discovered by
Kraushaar, uses small diffraction gratings to pro-
duce and recombine separate beams of light. The
usual two-parabolic-mirror schlieren system can
be converted inexpensively into a diffraction-grating
interferometer. Experimental data are presented
to verify the ability of the instrument to provide
valid and reliable measurements of air density.
Photographs of the flow in a supersonic cascade tun-
nel are included to indicate the quality of the inter-
NACA TN 2828
EFFECT OF A FINITE TRAILING-EDGE THICK-
NESS ON THE DRAG OF RECTANGULAR AND
DELTA WINGS AT SUPERSONIC SPEEDS. E. B.
Klunker and Conrad Rennemann, Jr. November
1952. 26p. diagrs. (NACA TN 2828)
The effect of a finite trailing-edge thickness on the
pressure drag of rectangular and delta wings with
truncated diamond-shaped airfoil sections with a
given thickness ratio is studied for supersonic Mach
numbers, linearized theory being used to evaluate
the surface pressures. In order to facilitate com-
parison with wings having sharp trailing edges, the
position of maximum thickness and base height are
determined for least pressure drag as functions of a
base-pressure parameter. Comparison is then
made between the drag of these wings and similar
wings with a sharp trailing edge for various aspect
ratios and thickness ratios as a iuncuon o01 stream
Mach number. The calculations of the drag
characteristics for these wings show that significant
drag reductions are possible under some conditions
at high supersonic speeds. These reductions are
relatively independent of aspect ratio for the rec-
tangular wings but depend considerably on aspect
ratio for the delta wings; the smaller aspect ratios
show the larger drag reductions. Calculations of
the spanwise distribution of drag are included to
compare further the effect of a base on the drag for
different aspect ratios.
NACA TN 2829
EXPERIMENTS ON TRANSONIC FLOW AROUND
WEDGES. George P. Wood. November 1952. 34p.
diagrs., photos., tab. (NACA TN 2829)
Several aspects of transonic flow around the forward
portions of wedge profiles were studied by means of
interferometry. Measurements were made of the
two kinds of flow pattern that occur at the leading
edge of a wedge at an angle of attack. The growth of
the supersonic region at a sharp convex corner
formed by two flat surfaces was observed. The
pressure-drag coefficients of a wedge of 14.5 semi-
angle were measured at Mach numbers of 0.768,
0.819, and 0.854 and were shown to be consistent
with those of wedges of smaller angle when plotted
according to the transonic similarity law. Conditions
at the bases of the shock waves that interacted with
boundary layers on the wedge were measured. The
method of characteristics was used to calculate the
flow behind an experimentally determined sonic line,
and the calculated flow field was compared with the
measured flow field. The accuracy in the location of
the sonic line necessary to give correctly the pres-
sure distribution on the surface behind it was deter-
NACA TN 2830
SEVERAL COMBINATION PROBES FOR SURVEYING
STATIC AND TOTAL PRESSURE AND FLOW DIREC-
TION. Wallace M. Schulze, George C. Ashby, Jr.
and John R. Erwin. November 1952. 64p. diagrs.,
photos., tab. (NACA TN 2830)
An investigation has been conducted to provide a
basis for design of combination probes intended to
survey static and total pressure and direction of flow
with special reference to subsonic turbomachine
testing. Static-pressure probes, yaw-element
probes, claw-type yaw probes, and combination
probes were tested in an 8-inch-diameter calibration
tunnel at air velocities up to 445 feet per second.
From the results of this investigation, the factors
which determine the sensitivity of claw-type yaw
probes were determined. Satisfactory combination
survey probes for sensing static and total pressure
and direction of flow in one or two planes were
NACA TN 2834
FLOW SURFACES IN ROTATING AXIAL-FLOW
PASSAGES. John D. Stanitz and Gaylord 0. Ellis.
November 1952. 31p. diagrs. (NACA TN 2834)
In order to investigate the deviation of flow surfaces
from their assumed orientation in the usual type of
two-dimensional solution, three-dimensional, in-
compressible, nonviscous, absolute irrotational
fluid motion is determined for flow through rotating
axial-flow passages bounded by straight blades of
finite spacing and infinite axial length lying on
meridional planes. Solutions are obtained for five
passages with varying blade spacing and hub-tip ratio.
The results are presented in such a manner as to
apply for all ratios of axial velocity to passage tip
speed. It is concluded that, for conditions in typical
axial-flow blade rows, the deviation of flow surfaces
RESEARCH ABSTRACTS NO.33
from their assumed orientation in two-dimensional
solutions is small.
NACA TN 2835
EFFECT OF CHANGING PASSAGE CONFIGURATION
ON INTERNAL-FLOW CHARACTERISTICS OF A
48-INCH CENTRIFUGAL COMPRESSOR. H -
CHANGE IN HUB SHAPE. John Mizisin and Donald
J. Michel. November 1952. 35p. diagrs., photo.,
tab. (NACA TN 2835)
The passage contour of a 48-inch centrifugal com-
pressor was modified by changing the shape of the
hub to control the deceleration rates along the blade
surfaces in order to improve the internal efficiency
of the impeller. A comparison of internal-flow char-
acteristics at design flow rate was made with the
original impeller and with a modified-blade impeller
that had the same area variation in the passage. In
addition, flow characteristics of the modified-hub
impeller over a flow range from maximum flow to
near surge at a corrected tip speed of 700 feet per
second are presented. At design flow, even though
the deceleration rate along the trailing face was less
for the modified-hub impeller than for the original
impeller, the clearance losses (increased by the
larger ratio of clearance to passage height) caused
the total-pressure losses (and thus the relative
adiabatic efficiency) to be about the same at the
impeller exit for the two configurations, thus pre-
cluding any improvement in over-all performance.
As in the original modified-blade impellers, large
losses occurred at the driving-face inlet at negative
angles of attack and in regions of large decelerations
along the trailing-face flow surface.
NACA RM 52125
TORSION, COMPRESSION, AND BENDING TESTS
OF TUBULAR SECTIONS MACHINED FROM 75S-T6
ROLLED ROUND ROD. R. L. Moore and J. W.
Clark, Aluminum Company of America. November
1952. 33p. diagrs., photos., 4 tabs. (NACA
Tests were made of tubular sections machined from
75S-T6 aluminum-alloy rolled rod and having ratios
of tube diameter to wall thickness D/t ranging from
2 to 150. The purpose of the investigation was to
establish curves of strength in torsion, compression,
and bending against D/t for the tubular sections and
to show to what extent these strengths may be cor-
related with the mechanical properties of the materi-
al. In view of the acceptable mechanical properties
obtained for the material, the relations obtained be-
tween these strengths and D/t may be used as a ten-
tative basis for design of members of the type inves-
NACA RM E52109
EFFECT OF RADIANT ENERGY ON VAPORIZA-
TION AND COMBUSTION OF LIQUID FUELS.
A. L. Berlad and R. R. Hibbard. November 1952.
46p. diagrs., 2 tabs. (NACA RM E52109)
RESEARCH ABSTRACTS NO.33
The radiative processes involved in combustion were
investigated to determine the present role of radiant
energy transfer in combustors. It was shown that at
present the amount of radiant energy transfer from
flame to fuel is quite small in a turbojet combustor.
In order to find techniques for making this radiant
energy transfer from flame to fuel significant, meth-
ods of increasing the equivalent gray-body emissivi-
ties of the fuel drops and the flame, as well as the
efficiency of the energy transfer itself, were ex-
amined. The equivalent gray-body emissivity of a
hydrocarbon fuel may be increased by the use of liq-
uid or solid, soluble or nonsoluble, additives. It
was found that the nonsoluble solid additive was the
most desirable one. In addition to the fact that a
slurry-type drop may have a greater emissivity, the
suspended solid additive may also greatly contribute
to the emissivity of the flame. An approximate equa-
tion was derived from which the equivalent gray-
body emissivities of such slurry-type drops may be
calculated. Data necessary for these calculations
were obtained by spectrophotometric analysis of thin
slurry films. In addition to the fuel additive, the use
of radiation-reflecting walls in a combustor was con-
sidered as a means of maintaining the efficiency of
the radiative transfer of energy from flame to fuel
drop. It was then shown that more than half the heat
of evaporation of the hydrocarbon constituent of this
slurry-type drop might be supplied by radiant energy
transfer in such a modified turbojet combustor.
NACA RM E52126
EXAMINATION OF SMOKE AND CARBON FROM
TURBOJET-ENGINE COMBUSTORS. Thomas P.
Clark. November 1952. 12p. photos., 2 tabs.
(NACA RM E52I26)
Smoke and carbon from turbojet-engine combustors
were studied by the methods of electron microscopy,
chemical analysis, and X-ray diffraction. The
smoke exhausting from a combustor was found to
consist of carbon black, agglomerated into soot.
The carbon black had been partially burned in its
passage through the flame zone. The smoke result-
ed from the incomplete combustion of the vaporized
fuel; it was not the result of the pyrolysis of fuel
droplets. The soft carbon in the dome of the com-
bustor liner was found to consist of carbon black and
soot intermixed with indeterminate complexes such
as high-boiling fuel ends and partly polymerized and
pyrolyzed heavy hydrocarbons. The hard carbon on
the walls of the combustor liner was found to be
largely a petroleum coke. The coke was apparently
formed by the liquid phase cracking, pyrolysis, and
subsequent coking on the liner wall of fuel from the
Forest Products Research Lab. (Gt. Brit.)
THE 'LEVER AID' IN THE KNIFE TEST OF PLY-
WOOD. P. Burgess. (Forest Products Research
Lab.; Reprint from Wood, v. 17, Sept., 1952,
This report describes a tool to aid in the conduction
of forcible separation or knife tests for assessing
the quality of the glue bond between the veneers of
Royal Aircraft Establishment (Gt. Brit.)
MEASUREMENT OF THE MOISTURE CONTENT OF
HIGH PRESSURE OXYGEN FOR USE IN AIRCRAFT.
II. A HIGH PRESSURE HYGROMETER. M. E.
Bedwell and W. G. Glendinning. May 1952. 18p.
diagrs., 3 tabs. (RAE Tech. Note Chem. 1175)
The gas flow bridge principle has been used in de-
signing an all-metal hygrometer to measure the
moisture content of oxygen gas at transport cylinder
pressure. The method could be applied to a produc-
ing plant, to give a warning if the moisture content of
the gas exceeded the maximum permitted in the spec-
ification. Both glass-capillary and metal-capillary
bridges have given satisfactory results at pressures
approaching that of the atmosphere. A magnetically
operated pressure sensitive device has been made to
function at the pressure-differential available, but
would not operate in the bridge because of slight gas-
leakage past the float, which upset the bridge off-
Royal Aircraft Establishment (Gt. Brit.)
THE DEVELOPMENT OF FREQUENCY AND VOLT-
AGE CONTROL SYSTEMS, EMPLOYING TRANS-
DUCTOR AMPLIFIERS, FOR A 1.75 KVA, 115 V,
1600 C/S MOTOR-ALTERNATOR, WITH SPECIAL
REFERENCE TO DYNAMIC CONSIDERATIONS.
C. W. Cooper, W. N. Corn and E. W. Eaton.
April 1952. 76p. diagrs., photos. (RAE EL. 1480)
This report describes the design, development, con-
struction and performance of Voltage and Frequency
control systems for a 1. 75 KVA, 115 V, 1600 c/s
Motor-Alternator intended for air-borne applications.
The control amplifiers employ transductors for pow-
er amplification. Details are given of the static and
dynamic performances of the amplifiers, and the
transient responses of the control systems.
Aeronautical Research Council (Gt. Brit.)
FREE STREAMLINE JETS IN SHEAR FLOW, AND
THEIR APPLICATION TO THE DESIGN OF SUCTION
SLOTS. W. T. Lord. June 20, 1950. 79p. diagrs.
(ARC 13, 197; FM 1450)
Solutions for the shapes of free streamline jets is-
suing from a gap between two flat parallel surfaces
may serve as a basis for designing slots for sucking
away the boundary layer on an airfoil. A brief
description is given of some unsuccessful attempts to
obtain analytical solutions for such jets when the flat
surfaces are collinear and the flow at a large distance
from the gap is a constant shearing motion containing
vorticity of any in.i'nirude A method of solution for
the case when the magnitude of the vorticity is suf-
ficiently small for terms of the order of its square
and higher powers to be neglected is given in detail.
This method may be extended to cover the more gen-
eral case when there is a step between the flat sur-
faces on each side of the gap.
Aeronautical Research Council (Gt. Brit.)
INCOMPIVESSIBLE TWO-DIMENSIONAL FLOW OF
AN INVISC [) FLUID ABOUT A SYMMETRICAL
AEROFOIL IN A CHANNEL OR FREE S6 BEAM
L. C. Woods. July 7, 1950. 32p. diagrs., 6 tabs.
(ARC 13,240; FM 1456)
The paper introduces a new and exact method of cal-
culating the incompressible, inviscid flow about a
symmetrical two-dimensional airfoil at zero inci-
dence. The asymmetric case and the effect of cir-
culation will be dealt with a subsequent paper. The
method gives velocities throughout the field and is
much quicker than the "relaxation" methods that have
been applied to this problem. The principle is to re-
place the airfoil by a series of small arcs on which
it is assumed that the product of the radius of curva-
ture and the velocity is constant. It is almost equiv-
alent to replacing the airfoil by a many sided polygon
for, at some distance from the airfoil, the field due
to a polygon and that due to a profile composed of
arcs as defined above, is sensibly the same. Each
arc, of course, must cover the same range as the
corresponding side of the polygon. The method is
termed "The Polygon Method. The idealization
takes place in the plane in which the airfoil is rep-
resented by a slit, that is, the flow plane. The
fields due to each of these arcs are added to give the
combined field due to the idealized airfoil. The
number of arcs selected to replace the airfoil is
governed by the accuracy required in the final re-
sults. Ii the typical example given at the end of the
paper, the author found that 12 arcs were sufficient
to give accuracies better than 1 percent in the ve-
locity increment due to the airfoil. In a later paper,
the method will be extended to obtain an approximate
solution of compressible flow, the accuracy of
which is a little better than that of the well-known
RESEARCH ABSTRACTS NO.33
Aeronautical Research Council (Gt. Brit.)
AN EXACT FOURIER SERIES METHOD OF CALCU-
LATIN, THE TWO-DIMENSIONAL INCOMPRESS-
IBLE FLOW ABOUT AN ASYMMETRICAL AERO-
FOIL. L. C. Woods. October 12, 1950. 6p.
diners. (ARC 13,440; FM 1484; Oxford Univ.,
LiEnginetring Lab. No. 46)
A method of t alc'ulalin the flow about an asymmetric
airfoil at incidence in a free stream is developed.
It involves the calculation of a number of Fourier
series, the constants of which rapidly converge to
zero. The flow about an asymmninietric airfoil at in-
cidence in a channel is tackled by an iterative pro-
cess which the author expects to be rapidly conver-
(Irtt Separate solutions are obtained in the two
regions separated by the stagnation streamlines and
are modified step by step until they match along
National Gas Turbine Establishment (Gt. Brit.)
THERMODYNAMIC PROPERTIES OF AIR AND
COMBUSTION PRODUCTS OF HYDROCARBON
FUELS, PART II. D. Fielding and J. E. C. Topps.
July 1952. 72p. diagrs., 7 tabs. (NGTE R. 120)
This report is a continuation of NGTE Report
No. R 74, civint methods whereby the data required
for gas turbine performance calculations, using any
fuel or supporter of combustion, may be rapidly
computed. The principle of the method given is the
addition of a calculated quantity to the corresponding
value of the property for dry air, so numerical con-
sistency is automatically ensured. The theory from
which the corrections are deduced is given fully, and
functions have been computed permitting determina-
tion of the true specific heat, the total heat and the
entropy function of the combustion products of all
common fuel-medium mixtures. The thermody-
namic aspects of the combustion process are exam-
ined and parameters for the calculation of combus-
tion temperature rise are deduced, and calculated
functions are given which permit this determination
for any fuel and any atmosphere. The data and
methods are suitable for the computation of working
data sheets, or for the direct determination of gas
turbine performance. Examples of the latter are
given, also nomographic aids. Special calculations
are discussed, and methods are developed whereby
data may be prepared for systems involving further
elements or compounds, combustible or otherwise,
in the fuel or atmosphere.
NACA Rept. 1050
Errata No. 1 on "FORMULAS FOR THE SUPER-
SONIC LOADING, LIFT AND DRAG OF FLAT
SWEPT-BACK WINGS WITH LEADING, EDGES BE-
HIND THE MACH LINES". Doris Cohen. 1951.
RESEARCH ABSTRACTS NO.33
DECLASSIFIED NACA REPORTS
NACA RM 8107
RESPONSE OF A ROTATING PROPELLER TO
AERODYNAMIC EXCITATION. Walter E. Arnoldi,
Hamilton Standard Propellers Division, United Air-
craft Corporation. January 21, 1949. 26p. diagrs.
(NACA RM 8107) (Declassified from Restricted,
This report presents a theoretical analysis of the
flexural vibration of an aircraft propeller blade sub-
jected to harmonic aerodynamic exciting forces at a
fixed multiple of propeller rotation frequency. The
flexural vibration of a rotating propeller blade with
clamped shank is analyzed with the object of present-
ing, in matrix form, equations for the elastic bend-
ing moments in forced vibration resulting from aero-
dynamic forces applied at a fixed multiple of rota-
tional speed. Matrix equations are also derived
which define the critical speeds and mode shapes for
any excitation order and the relation between critical
speed and blade angle.
NACA RM E7H21
PRELIMINARY I N'VEST ILuATION OF HYDRAZINE AS
A ROCKET FUEL. Paul M. Ordin, Riley 0. Miller
and John M. Diehl. May 24, 1948. 35p. diagrs.,
photos. (NACA RM E7H21) (Declassified from
Several properties of hydrazine were investigated.
Theoretical calculations were made of performance
of hydrazine with liquid 03, liquid 02, H202, and
HN03. Hydrazine is poisonous, decomposes some
materials, is temperature sensitive, does not propa-
gate detonation waves, reacts violently with H202,
NaMn04, C(N02)4, and HN03- N02, and gives high
theoretical performance with several oxidants.
NACA RM L6Ll1
WIND-TUNNEL INVESTIGATION OF WING INLETS
FOR A FOUR-ENGINE AIRPLANE. Walter A.
Bartlett, Jr. and Edwin B. Goral. March 11, 1947.
65p. diagrs., photos., 7 tabs. (NACA RM L6Lll)
(Declassified from Restricted, 9/16/52)
An investigation has been conducted in the Langley
propeller-research tunnel to develop wing-leading-
edge inlets for location between the inboard and out-
board nacelles on each wing of a four-engine airplane
for the Army Air Forces. The investigation includ-
ed aerodynamic tests of the basic wing and the origi-
nal inlet, and the development by the NACA of wing
inlets for two versions of the airplane. The original
inlet was found to decrease the maximum lift coef-
ficients and to have critical Mach numbers below
those of the wing with the basic nose installed. The
total-pressure recovery in the oil-cooler ducts was
poor regardless of the inlet installation. As the
sharp expanding bend in this duct cannot be avoided,
it is recommended that the oil-cooler air be induced
through the cowling or from some source other than
the subject wing inlet. Two inlets were developed
that should be satisfactory for the airplane. The
maximum lift coefficients for the model with inlets
5 and 6 installed were about 1. 21 and 1. 22, respect-
ively, with 0 wing flaps and 1. 87 and 2. 00, respect-
ively, with 65 wing flaps compared to correspond-
ing values of 1. 20 and 2. 01 for the model equipped
with the faired basic nose. The predicted critical
Mach numbers for inlets 5 and 6 for the critical mil-
itary power high-speed condition for an altitude of
40, 000 feet were 0. 63 and 0. 64, respectively, as
compared to 0. 64 for the thickest section of the basic
NACA RM L7E20
CORRELATION OF WIND-TUNNEL AND FLIGHT
DETERMINATIONS OF THE BUFFET SPEED OF AN
AIRPLANE EQUIPPED WITH EXTERNAL STORES.
H. Norman Silvers and Kenneth P. Spreemann.
March 2, 1948. 54p. diagrs., photos., tab. (NACA
RM L7E20) (Declassified from Restricted, 9/16/52)
Presents a comparison of wind-tunnel measurements
of incremental lift, drag, and pitching-moment coef-
ficients due to external stores and measurements of
pressures in the vicinity of the external stores with
flight determinations of buffet Mach numbers. It is
indicated that the variation of the incremental drag
coefficient with Mach number of the external store
may be used to estimate maximum flight buffet Mach
number, and the attainment of local critical pressure
in the vicinity of the store is indicative of the lower
limit of flight buffet due to stores.
NACA RM L7K10
WIND-TUNNEL INVESTIGATION OF THE STABIL-
ITY OF JETTISONED NOSE SECTIONS OF THE
D-558 AIRPLANE PHASES I AND II. Stanley
H. Scher. January 14, 1948. 33p. photos.,
diagrs., 6 tabs. (NACA RM L7K10) (Declassified
from Confidential, 9/16/52)
An investigation of the stability of models of the
jettisonable nose sections of the D-558 airplanes,
Phases I and II, has been conducted in the Langley
20-foot free-spinning tunnel. It was found that both
model nose sections tumbled end over end about an
approximately horizontal axis. The installation of
suitable fins, together with sufficient forward loca-
tion of the center of gravity, prevented the tumbling
and caused the models to descend in a stable nose-
NACA RM L8E11
PRELIMINARY INVESTIGATION TO DETERMINE
PROPELLER SECTION CHARACTERISTICS BY
MEASURING THE PRESSURE DISTRIBUTION ON AN
NACA 10-(3)(08)-03 PROPELLER UNDER OPERAT-
ING CONDITIONS. Albert J. Evans and George
Liner. July 14, 1948. 52p. Jiarrs., tab. (NACA
RM L8Ell) (Declassified from Confidential,
Contains preliminary data obtained by measuring the
pressure distribution on the airfoil sections at nine
radial stations of an NACA 10-(3)(08)-03 design pro-
peller. The results are presented as normal-force
and moment coefficients for a range of nominal
angle of attack (simple blade element theory) from 0
to 4 for a section Mach number range of about 0. 6
to 1. 15 for the outboard stations and about 0. 3 to 0. 6
for inboard stations.
NACA RM L8H11
AERODYNAMIC CHARACTERISTICS OF FLYING-
BOAT HULLS HAVING LENGTH-BEAM RATIOS OF
20 AND 30. John M. Riebe. November 10, 1948.
26p. diagrs., photos., 4 tabs. (NACA RM L8H11I)
(Declassified from Restricted, 9/16/52)
An investigation of a series of hulls of length-beam
ratios 6 to 15 previously reported in NACA TN 1305
has been extended to length-beam ratios 20 and 30.
The hulls of the entire series were designed to have
approximately the same hydrodynamic performance
with respect to spray and resistance characteristics
regardless of length-beam ratio. Drag coefficients
and longitudinal-stability parameters are presented
with and without wing interference. Lateral-stability
parameters are presented with wing interference.
NACA RM L9G05
EFFECT OF AN INCREASE IN HULL LENGTH-
BEAM RATIO FROM 15 TO 20 ON THE HYDRODY-
NAMIC CHARACTERISTICS OF FLYING BOATS.
Arthur W. Carter and Walter E. Whitaker, Jr.
August 24, 1949. 29p. diagrs., photos., 2 tabs.
(NACA RM L9G05) (Declassified from Restricted,
Investigations of the effect of hull lenth-beam ratio
on the hydrodynamic qualities of flying boats, made
in Langley tank No. 1, have been extended to include
a length-beam ratio of 20. The hydrodynamic qual-
ities determined in this investigation were longitu-
dinal stability during take-off and landing, spray
characteristics, and take-off performance in smooth
water and take-off and landing behavior in waves.
Massachusetts Inst. of Tech., Research Lab. for
Mechanics of Materials.
AN INVESTIGATION OF THE TORSION CREEP-TO-
RUPTURE PROPERTIES OF ALLOY N-155. M. E.
Shank. (Final rept.) April 4, 1952. 34p. diagrs.,
photos. (Massachusetts Inst. of Tech., Research
Lab. for Mechanics of Materials)
This report describes an investigation of the torsion
creep-to-rupture properties of alloy N-155. All
tests were in the range of 1200 to 1500 F, with
times in the 10-1000 hour range. The effect of
stress concentration in the form of transverse cir-
cular holes was investigated under similar condi-
tions. The effect of specimen geometry, as affected
by slii.ht changes in specimen contours, was found
UNIVERSITY OF FLORIDA
fll llhl f 1 111111 11111111111111111 111111 llll fll / 11111111
3 1262 09079 7464
RESEARCH ABSTRACTS NO.33
to be great. Preliminary investigation was made in-
to what appeared to be axial extension of the speci-
men during torsional creep.
APPLICATION OF TRANSONIC SIMILARITY CON-
SIDERATIONS TO FLOW ALONG A WAVE-SHAPED
WALL. Carl Kaplan. 1952. 42p. diagrs., tab.
(Presented at Eighth International Congress on Theo-
retical and Applied Mechanics held in Istanbul,
Turkey, August 20-28, 1952).
The Prandtl-Busemann small-perturbation method is
utilized to obtain the flow of a compressible fluid
past an infinitely long wave-shaped wall. When the
essential assumption for transonic flow (that all Mach
numbers in the region of flow are nearly unity) is In-
troduced, the expression for the velocity potential
takes the form of a power series in the transonic
similarity parameter. On the basis of this form of
the solution, an attempt is made to solve the non-
linear differential equation for transonic flow past a
wavy wall. The analysis utilized exhibits clearly
the difficulties inherent in nonlinear-flow problems.
Nevertheless, by means of the classical method of
Integration in series, recursion formulas for certain
four-labelled coefficients are obtained. Remarkably
enough, these recursion formulas can be solved
analytically and lead to a number of conclusions; for
example, that the transonic similarity parameter
cannot be greater than four-thirds. In addition, a
novel numerical test of convergence, applied to the
power series (in the transonic similarity parameter)
representing the local Mach number distribution at
the boundary, indicates that smooth symmetrical
potential flow past the wavy wall is no longer possi-
ble once the critical value of the undisturbed-stream
Mach number has been exceeded.
NACA-Langley 12-3-52 4M
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