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:
AA00009235:00088

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National Advisory Committee for Aeronautics


Research Abstracts


N0.27


AUGUST 15,1952


CURRENT NACA REPORTS


NACA Rept. 1025

EXPERIMENTAL AND THEORETICAL STUDIES OF
AREA SUCTION FOR THE CONTROL OF THE
LAMINAR BOUNDARY LAYER ON AN NACA 64A010
AIRFOIL. Albert L. Braslow, Dale L. Burrows,
Neal Tetervin and Fioravante Visconti. 1951. ii,
19p. diagrs., photos. (NACA Rept. 1025. Formerly
TN 1905; TN 2112)

Low-turbulence wind-tunnel tests were made of an
NACA 64A010 airfoil section having a porous surface
to investigate continuous suction as a means of con-
trolling the laminar boundary layer. Full-chord lam-
inar flow and large net reductions in drag were ob-
tained up to a Reynolds number of approximately
20 x 10. It seems likely from the experimental re-
sults and a related theoretical analysis that attain-
ment of full-chord laminar flow by means of continu-
ous suction through a porous surface will be possible
at any value of the Reynolds number provided that the
airfoil surfaces are maintained sufficiently smooth
and fair and provided that outflow of air through the
surface is prevented.


NACA Rept. 1029

COMPRESSIVE STRENGTH OF FLANGES. Elbridge
Z. Stowell. 1951. 14p. diagrs., tab. (NACA
Rept. 1029. Formerly TN 2020)

The maximum compressive stress carried by a
hinged flange is computed from a deformation theory
of plasticity combined with the theory for finite de-
flections for this structure. The computed stresses
agree well with those found experimentally. Empiri-
cal observation indicates that the results will also
apply fairly well to the more commonly used flanges
which are not hinged.

NACA Rept. 1030

INVESTIGATION OF SEPARATION OF THE TURBU-
LENT BOUNDARY LAYER. G. B. Schubauer and
P. S. Klebanoff, National Bureau of Standards. 1951.
20p. diagrs., photos., 8 tabs. (NACA Rept. 1030.
Formerly TN 2133)

An investigation was conducted on a turbulent bound-
ary layer near a smooth surface with pressure
gradients sufficient to cause flow separation. The
Reynolds number was high, but speeds were entirely
within the incompressible flow range. The investiga-
tion consisted of measurements of mean flow, three
components of turbulence intensity, turbulent shear-


ing A-.tres.-, and correlati-$ns between two fluctuation
components at a point and between the same conmp-i-
nenk!t dilterent point.. Results are given in the
form ol tables, and graphs.. The discussion deal-.
first with separation and then with the more funda-
mental question of basic concepts of turbulent flow.


NACA Rept. 1035

ANALYSIS OF MEANS OF IMPROVING THE UN-
CONTROLLED LATERAL MOTIONS OF PERSONAL
AIRPLANES. Marion 0. McKinney, Jr. 1951. ii,
9p. diagrs., 3 tabs. (NACA Rept. 1035. Formerly
TN 1997)

A theoretical analysis of means of improving the un-
controlled lateral motions of personal airplanes is
presented. The purpose of this analysis was to deter-
mine whether such airplanes could be made capable
of flying uncontrolled for an indefinite period of
time without getting into a dangerous attitude and for
a reasonable period of time without deviating ex-
cessively from the original course.


NACA Rept. 1037

GENERAL METHOD AND THERMODYNAMIC
TABLES FOR COMPUTATION OF EQUILIBRIUM
COMPOSITION AND TEMPERATURE OF CHEMICAL
REACTIONS. Vearl N. Huff, Sanford Gordon and
Virginia E. Morrell. 1951. ii, 57p., 45 tabs.
(NACA Rept. 1037. Formerly NACA TN 2113,
TN 2161)

A rapidly convergent successive approximation
process is described that simultaneously determines
both composition and temperature resulting from a
chemical reaction. This method is suitable for use
with any set of reactants over the complete range of
mixture ratios as long as the products of reaction are
ideal gases. An approximate treatment of limited
amounts of liquids and solids is also included. This
method is particularly suited to problems having a
large number of products of reaction and to problems
that require determination of such properties as
specific heat or velocity of sound of a dissociating
mixture. The method presented is applicable to a
wide variety of problems that include (1) combustion
at constant pressure or volume; and (2) isentropic
expansion to an assigned pressure, temperature, or
Mach number. Tables of thermodynamic functions
needed with this method are included for 42 sub-
stances for convenience in numerical computations.


*AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASHINGTON 25, D. C, CITING CODE NUMBER ABOVE EACH TITLE,
THE REPORT TITLE AND AUTHOR.







2

NACA Rept. 1038

WIND-TUNNEL INVESTIGATION OF AIR INLET
AND OUTLET OPENINGS ON A STREAMLINE
BODY. John V. Becker. 1951. ii, 21p. diagrs.,
photos., 3 tabs. (NACA Rept. 1038. Formerly
ACR, November 1940)

Air inlet openings at the nose of a streamline body
and outlets at the tail and at the 21-percent and 63-
percent stations of the body were investigated with
the objective of developing low-drag high-critical-
speed configurations. Drag, pressure-distribution,
and boundary-layer measurements were made. Inlet
and outlet opening shapes were found which caused no
increase in drag over that of the basic streamline
body to which the openings were applied. The criti-
cal Mach number of the ducted body was also equal to
that of the basic body over a wide range of internal
air flow rates.


NACA Rept. 1042

SOME EFFECTS OF NONLINEAR VARIATION IN
THE DIRECTIONAL-STABILITY AND DAMPING-IN-
YAWING DERIVATIVES ON THE LATERAL STA-
BILITY OF AN AIRPLANE. Leonard Sternfield.
1951. ii, 9p. diagrs., tab. (NACA Rept. 1042.
Formerly TN 2233)

The effect of nonlinear stability derivatives on the
lateral stability of an airplane is analyzed. Motions
are calculated on the assumption that the values of
the directional-stability derivative and the damping-
in-yawing derivative are functions of the angle of
sideslip. The application of the Laplace transform to
the calculation of an airplane motion when certain
types of nonlinear derivatives are present is de-
scribed in detail.


NACA Rept. 1047

THE STABILITY OF THE COMPRESSION COVER OF
BOX BEAMS STIFFENED BY POSTS. Paul Seide
and Paul F. Barrett. 1951. ii, 16p. diagrs., 3 tabs.
(NACA Rept. 1047. Formerly TN 2153)

An investigation is made of the buckling of the com-
pression cover of post-stiffened box beams subjected
to end moments. Charts are presented for the deter-
mination of the minimum post axial stiffnesses and
the corresponding compressive buckling loads re-
quired for the compression cover to buckle with nodes
through the posts. Application of the charts to de-
sign and analysis and the limitations of their use are
discussed.


NACA Rept. 1049

EXPERIMENTAL INVESTIGATION OF THE EFFECT
OF VERTICAL-TAIL SIZE AND LENGTH AND OF
FUSELAGE SHAPE AND LENGTH ON THE STATIC
LATERAL STABILITY CHARACTERISTICS OF A
MODEL WITH 45 SWEPTBACK WING AND TAIL
SURFACES. M. J. Queijo and Walter D. Wolhart.
1951. ii, 29p. diagrs., photos., 4 tabs.
(NACA Rept. 1049. Formerly TN 2168)


NACA
RESEARCH ABSTRACTS NO.27


Results are presented of an investigation to deter-
mine the effects of vertical-tail size and length and
of fuselage shape and length on the lateral-static
stability derivatives of a midwing airplane model
with 45 sweptback wing and tail surfaces. Inter-
ference between the various model components also
is evaluated.


NACA Rept. 1052

A SUMMARY OF LATERAL-STABILITY DERIVA-
TIVES CALCULATED FOR WING PLAN FORMS IN
SUPERSONIC FLOW. Arthur L. Jones and Alberta
Alksne. 1951. ii, 35p. diagrs., 3 tabs. (NACA
Rept. 1052)

Values of the lateral-stability derivatives for wings
at supersonic speeds, calculated using the linearized
theory for compressible flow, are presented in the
form of design charts showing the variations of the
derivatives with Mach number and aspect ratio for
six plan forms. Limitations in the applicability and
availability of the lateral-stability derivatives are
discussed.


NACA Rept. 1059

A BIHARMONIC RELAXATION METHOD FOR CAL-
CULATING THERMAL STRESS IN COOLED IRREGU-
ULAR CYLINDERS. Arthur G. Holms. 1952. ii,
19p. diagrs., 4 tabs. (NACA Rept. 1059. Formerly
TN 2434)

A numerical method was developed for calculating
thermal stresses in irregular cylinders cooled by one
or more internal passages. The use of relaxation
methods and elementary methods of finite differences
was found to give approximations to the correct values
when compared with previously known solutions for
concentric circular cylinders possessing symmetrical
and asymmetrical temperature distributions.


NACA Rept. 1060

DETAILED COMPUTATIONAL PROCEDURE FOR
DESIGN OF CASCADE BLADES WITH PRESCRIBED
VELOCITY DISTRIBUTIONS IN COMPRESSIBLE PO-
TENTIAL FLOWS. George R. Costello, Robert L.
Cummings and John T. Sinnette, Jr. 1952. ii, 14p.
diagrs., 9 tabs. (NACA Rept. 1060. Formerly
TN 2281)

A detailed step-by-step computational outline is pre-
sented for the design of cascade blades having a pre-
scribed velocity distribution on the blade in a poten-
tial flow of the usual compressible fluid. This outline
is based on the assumption that the magnitude of the
velocity in the flow of the usual compressible non-
viscous fluid is proportional to the magnitude of the
velocity in the flow of a compressible nonviscous fluid
with linear pressure-volume relation. The computa-
tional procedure includes several ways of adjusting
the prescribed velocity to satisfy the restriction im-
posed by the method. Tables of coefficients are
given for evaluating the necessary integrals, includ-
ing the determination of the harmonic conjugate. Nu-
merical examples are included.






NACA
RESEARCH ABSTRACTS N0.27


NACA Rept. 1061

EFFECT OF INITIAL MIXTURE TEMPERATURE ON
FLAME SPEED OF METHANE-AIR, PROPANE-
AIR AND ETHYLENE-AIR MIXTURES. Gordon L.
Dugger. 1952. ii, 12p. diagrs., photo., 3 tabs.
(NACA Rept. 1061. Formerly TN 2170; TN 2374)

Flame speeds based on the outer edge of the shadow
cast by the laminar Bunsen cone were determined as
functions of composition for methane-air mixtures at
initial mixture temperatures ranging from -132 to
342 C and for propane-air and ethylene-air mixtures
at initial mixture temperatures ranging from -73 to
3440 C. The data showed that maximum flame speed
increased with temperature at an increasing rate.
The percentage change in flame speed with change in
initial temperature for the three fuels followed the de-
creasing order, methane, propane, and ethylene.
Empirical equations were determined for maximum
flame speed as a function of initial temperature over
the temperature range covered for each fuel. For
each fuel it was found that, with a fixed parallel-beam
shadowgraph system, the ratio of flame speed based
on the outer edge of the shadow cast by the flame cone
to flame speed based on the inner edge of the shadow
was a constant, independent of temperature or com-
position. The flame speed of propane-air flames was
independent of tube diameter from 10 to 22 milli-
meters or stream-flow Reynolds number from 1500 to
2100. The observed effect of temperature on flame
speed for each of the fuels was reasonably well pre-
dicted by either the thermal theory as presented by
Semenov or the square-root law of Tanford and Pease.
The importance of active radicals in flame propaga-
tion was indicated by a simple linear relation between
maximum flame speed and equilibrium radical con-
centrations for all three fuels.


NACA TN 2643

SPAN LOAD DISTRIBUTIONS RESULTING FROM
ANGLE OF ATTACK, ROLLING, AND PITCHING
FOR TAPERED SWEPTBACK WINGS WITH STREAM-
WISE TIPS. SUPERSONIC LEADING AND TRAILING
EDGES. John C. Martin and Isabella Jeffreys. July
1952. 143p. diagrs., 6 tabs. (NACA TN 2643)

On the basis of the linearized supersonic-flow theory
the span load distributions resulting from constant
angle of attack, from steady rolling, and from steady
pitching were calculated for a series of thin swept-
back tapered wings with streamwise tips and with
supersonic leading and trailing edges. The results
are valid for the Mach number range for which the
Mach line from either wing tip does not intersect the
remote half-wing. The results of the analysis are
presented as a series of design charts. Some illus-
trative variations of the spanwise distributions of
circulation with the various design parameters are
also presented.


NACA TN 2656

A BLADE-ELEMENT ANALYSIS FOR LIFTING
ROTORS THAT IS APPLICABLE FOR LARGE IN-
FLOW AND BLADE ANGLES AND ANY REASON-


3


ABLE BLADE GEOMETRY. Walter Castles, Jr.
and Noah C. New, Georgia Institute of Technology.
July 1952. 63p. diagrs., 7 tabs. (NACA TN 2656)

A blade-element analysis for lifting rotors is pre-
sented which is applicable for large inflow and blade
angles and any reasonable blade geometry and should
therefore be useful for convertaplane as well as heli-
copter calculations. Simple approximate relations
between rotor thrust and flight-path velocity compo-
nents and rotor blade angles, torque, and in-plane
forces are derived and these solutions, based on the
assumption of a triangular distribution of blade
circulation and a parabolic variation of profile drag
with lift, are sufficiently accurate for preliminary
calculations and determination of equilibrium angle of
attack and lateral tilt of the tip-path plane. More
exact blade-element equations are then derived for
the relations between thrust and flight-path velocity
components and equilibrium blade angles, torque,
and in-plane forces and moments.


NACA TN 2675

MEASUREMENTS OF FLYING QUALITIES OF AN
F-47D-30 AIRPLANE TO DETERMINE LATERAL
AND DIRECTIONAL STABILITY AND CONTROL
CHARACTERISTICS. R. Fabian Goranson and
Christopher C. Kraft, Jr. July 1952. 61p. diagrs.,
photos., 2 tabs. (NACA TN 2675)

Tests have been made of the flying qualities of an
F-47D-30 airplane to determine the lateral and
directional stability and control characteristics.
Data are also presented of the aileron hinge-moment
characteristics.


NACA TN 2715

THE THEORETICAL CHARACTERISTICS OF
TRIANGULAR-TIP CONTROL SURFACES AT SUPER-
SONIC SPEEDS. MACH LINES BEHIND TRAILING
EDGES. Julian H. Kainer and Mary Dowd King.
July 1952. 76p. diagrs., 4 tabs. (NACA TN 2715)

By means of linearized theory, generalized expres-
sions in closed form have been obtained for the char-
acteristics due to control-surface deflection
(CL6 C16, Cm6, and Ch6 ) and due to wing angle
of attack (Cha) for wing plan forms having triangular-
tip control surfaces at supersonic speeds. The
analysis considers wing trailing-edge sweep, control-
surface geometry, and Mach number for the deflec-
tion characteristics. For Cha, the effects of wing

leading-edge sweep and aspect ratio are also included.
The analysis is limited to configurations where the
trailing edges are supersonic and where the inner-
most Mach lines from the leading edge of the control-
surface root chord do not intersect the wing root
chord.








4


NACA TN 2737

PLASTIC STRESS-STRAIN RELATIONS FOR COM-
BINED TENSION AND COMPRESSION. Joseph
Marin and H. A. B. Wiseman, Pennsylvania State
College. July 1952. 61p. diagrs., photos., 2 tabs.
(NACA TN 2737)

I'lastic stress-strain relations for biaxial tension-
compression principal stresses were determined for
a 14S-T4 aluminum alloy. Constant-stress-ratio
tests provided control data and information on the in-
fluence of biaxial stresses on the yield strength of
the material. Variable-stress-ratio tests were made
to determine whether the deformation or flow theory
a.ii ee better with actual plastic stress-strain rela-
tions and special tests were conducted to check the
validity of various assumptions made in these
theories.


NACA TN 2739

NUMERICAL DETERMINATION OF INDICIAL LIFT
AND MOMENT FUNCTIONS FORA TWO-
DIM ENSIONAL SINKING AND PITCHING AIRFOIL AT
MACH NUMBERS 0.5 AND 0.6. Bernard Mazelsky
and Joseph A. Drischler. July 1952. 37p. diagrs.,
4 tabs. (NACA TN 2739)

Approximate values for the indicial lift and moment
functions on a sinking and pitching airfoil are calcu-
lated for Mach numbers 0.5 and 0.6. The indicial
lift function associated with an airfoil penetrating a
sharp-edge gust is also determined approximately at
Mach numbers 0.5, 0.6, and 0.7.


NACA TN 2741

INVESTIGATION OF THE INFLUENCE OF FUSE-
LAGE AND TAIL SURFACES ON LOW-SPEED
STATIC STABILITY AND ROLLING CHARACTERIS-
TICS OF A SWEPT-WING MODEL. John D. Bird,
Jacob H. Lichtenstein and Byron M. Jaquet. July
1952. 18p. diagrs., photo. (NACA TN 2741. For-
merly RM L7H15)

Results are presented of a wind-tunnel investigation
to determine influence of the fuselage and tail on
static stability and rotary derivatives in roll of a
model having 45 sweptback wing and tail surfaces.
The wing alone and the model without the horizontal
tail showed marginal longitudinal stability near maxi-
mum lift. The longitudinal stability of the complete
model was satisfactory. The vertical tail produced
larger increments of rate of change of lateral-force
and yawing-moment coefficients with wing-tip helix
angle than the fuselage or the horizontal tail.


NACA TN 2744

PRACTICAL CALCULATION OF SECOND-ORDER
SUPERSONIC FLOW PAST NONLIFTING BODIES OF
REVOLUTION. Milton D. Van Dyke. July 1952.
62p. diagrs., 2 tabs., 2 charts. (NACA TN 2744)


NACA
RESEARCH ABSTRACTS NO. 27


Calculation of second-order supersonic flow past
bodies of revolution at zero angle of attack is de-
scribed in detail, and reduced to routine computation.
Use of an approximate tangency condition is shown to
increase the accuracy for bodies with corners.
Tables of basic functions and standard computing
forms are presented. The procedure is summarized
so that one can apply it without necessarily under-
standing the details of the theory. A sample calcula-
tion is given, and several examples are compared
with solutions calculated by the method of character-
istics.



NACA TN 2745

INFLUENCE OF CHEMICAL COMPOSITION ON
RUPTURE TEST PROPERTIES AT 1500 F OF
FORGED CHROMIUM-COBALT- NICKEL-IRON BASE
ALLOYS. J. W. Freeman, J. F. Ewing and A. E.
White, University of Michigan. July 1952. 69p.
diagrs., photos., 2 tabs. (NACA TN 2745)

Six1iy-two alloys involving systematic individual varia-
tions of 10 elements present in a forged chromium-
cobalt-nickel-iron base alloy in the solution-treated
and aged condition and simultaneous variation of
molybdenum, tungsten, and columbium were rupture
tested at 1500 F. It was found that the elements can
be varied individually between quite wide limits with-
out significantly changing the rupture properties. The
results were similar to those previously obtained at
1200 F (NACA Rept. 1058) except that increased
chromium was not beneficial at 1500 F and the
saturation effect for molybdenum and tungsten did not
occur at 1200 F.


NACA TN 2746

PREVIEW OF BEHAVIOR OF GRAIN BOUNDARIES
IN CREEP OF ALUMINUM BICRYSTALS. F. N.
Rhines and A. W. Cochardt, Carnegie Institute of
Technology. July 1952. 40p. diagrs., photos.
(NACA TN 2746)

Gliding of aluminum bicrystals along their mutual
boundary was studied during creep tests at 200 to
650 C and 1 to 100 psi. The direction of motion de-
pends only upon the direction of maximum resolved
shear stress in the boundary plane. The magnitude of
the motion increases with magnitude of resolved shear
stress, temperature, and degree of mismatch between
orientations of the crystals. Gliding rate is cyclic
and for each temperature there is a stress below
which gliding does not occur. At low temperature and
small orientation difference an induction period pre-
cedes gliding. Gliding operates in a relatively thick
zone of crystalline metal on both sides of the bound-
ary. Portions of the grains away from the boundary
degenerate into blocks that move with respect to each
other, parallel to the tension axis, and that rotate
about octahedral axes of the original crystal.







NACA
RESEARCH ABSTRACTS NO.27


NACA TN 2750

MATRIX AND RELAXATION SOLUTIONS THAT
DETERMINE SUBSONIC THROUGH FLOW IN AN
AXIAL-FLOW GAS TURBINE. Chung-Hua Wu.
July 1952. 65p. diagrs., 7 tabs. (NACA TN 2750)

A method recently developed for determining the
steady flow of a nonviscous compressible fluid along
a relative stream surface between adjacent blades in
a turbomachine was applied to investigate subsonic
through flow in a single-stage axial-flow gas turbine.
For both incompressible and compressible flows, the
variation in flow on the stream surface was deter-
mined by using both the relaxation and the matrix
methods. In all solutions considered, convergence
was obtained without difficulty. The compressibility
of the gas and the radial twist of the stream surface
had equally significant effects on the considerable
amount of radial flow obtained. The results of these
accurate calculations provide a basis for evaluation
of simpler, more approximate methods for computing
subsonic through flow in turbomachines.



NACA TN 2754

A METHOD OF SELECTING THE THICKNESS,
HOLLOWNESS, AND SIZE OF A SUPERSONIC WING
FOR LEAST DRAG AND SUFFICIENT BENDING
STRENGTH AT SPECIFIED FLIGHT CONDITIONS.
James L. Amick. July 1952. 38p. diagrs. (NACA
TN 2754)

The method described, which is suitable for use in
preliminary missile design, is based on simple beam
theory and takes into account the effects of wing
weight and type of skin thickness distribution. For
cases in which the wing weight is negligible, a
simplification of the general method gives the best
lift-drag ratio consistent with the bending strength
requirement of a given combination of plan form and
profile shape at a given Mach number as a function of
a single design parameter. An example of the appli-
cation of the method to a diamond wing at Mach num-
ber 2.0, for a range of specified flight conditions, is
presented.


NACA RM E52C31

TRUE AIRSPEED MEASUREMENT BY IONIZATION-
TRACER TECHNIQUE. Bemrose Boyd, Robert G.
Dorsch and George H. Brodie. July 1952. 37p.
diagrs., photos. (NACA RM E52C31)

Ion bundles produced in a pulse-excited corona dis-
charge are used as tracers with a radar-like pulse
transit-time measuring instrument in order to pro-
vide a measurement of airspeed that is independent of
all variables except time and distance. The resulting
instrumentation need not project into the air stream
and, therefore, will not cause any interference in
supersonic flow. The instrument was tested at Mach
numbers ranging from 0.3 to 3.8. Use of the proper
instrumentation and technique results in accuracy of
the order of 1 percent.


5


NACA RM E52E12

USE OF CHOKED NOZZLE TECHNIQUE AND
EXHAUST JET DIFFUSER FOR EXTENDING OPER-
ABLE RANGE OF JET-ENGINE RESEARCH FACIL-
ITIES. John H. Povolny. July 1952. 17p. diagrs.
(NACA RM E52E12)

An investigation has been conducted to determine the
increase in the useful ranges of flight conditions that
may be obtained with a given jet-engine research
facility when the choked nozzle technique or the
exhaust jet diffuser, or both, are employed. This
report describes these two methods, presents the
considerations involved in their application, and
gives typical results of their use as well as confir-
mation of the accuracy of data obtained by utilization
of these techniques. The validity and accuracy of the
choked nozzle technique and the associated area -
pressure-differential thrust correction term were
substantiated by turbojet-engine and exhaust-nozzle
performance data covering a range of nozzle pres-
sure ratios up to about 10. It was demonstrated by
calculations for a typical turbojet engine installed in
a typical altitude test facility that a considerable in-
crease in the range of flight conditions that can be in-
vestigated may be obtained by use of the choked
nozzle technique. It was also demonstrated that the
range of facility exhaust pressures or exhauster
flows may be increased by use of the exhaust jet
diffuser.


BRITISH REPORTS


N-16326*

Aeronautical Research Council (Gt. Brit.)
PUBLISHED REPORTS AND MEMORANDA OF THE
AERONAUTICAL RESEARCH COUNCIL. 1952. 5p.
(ARC R & M 2350)

Lists R & M's 2251 through 2350 and gives the Coun-
cil number, the title and the author for each.


N-16327*

Aeronautical Research Council (Gt. Brit.)
TUNNEL BLOCKAGE NEAR THE CHOKING CONDI-
TION. A. Thorn and Myra Jones. 1952. 16p.
diagrs., 2 tabs. (ARC R & M 2385; ARC 8606;
ARC 8878. Formerly RAE Aero 2020; RAE
Aero 2056; ARC 9854; FM 968)


By a consideration of mass flow, blockage correc-
tions are estimated up to the choking speed of the tun-
nel, and a curve is given showing the blockage for a
Rankine oval in terms of the uncorrected choking
Mach number. The theory of the method of deter-
mining blockage from wall pressures is examined
more critically for the two-dimensional cases and
extended to certain three-dimensional cases.







6


N-16328*

Aeronautical Research Council (Gt. Brit.)
BOUND AND TRAILING VORTICES IN THE LINE-
ARISED TH Oi-'Y OF SUPERSONIC FLOW, AND THE
DOWNWASH IN THE WAKE OF A DELTA WING.
A. Robinson and J. H. Hunter-Tod. 1952. 14p.
dij.rs. (ARC R :. M 2409; ARC 11,296. Formerly
College of Aeronautics, Crfijiilid Rept. 10)

The field of flow round a flat airfoil at incidence can
be regarded in linearized theory as the result of both
bound and trailing vortices for supersonic as well as
for low-speed flight. This leads to a convenient
method, given the lift distribution over an iirfil. for
( .ili ulailinL the flow round it at supersonic speeds.
As an application of the results, the downwash is cal-
culated in the wake of a delta wing lying within the
Mach cone .iimnatin4 from its apex. The downwash
is found to be least just aft the trailing edge and is
everywhere less than the downflow at the airfoil. It
increases steadily to a lniiliii value which is attain-
ed virtually within two chord lengths of the trailing
edil-'t The ratio of the downwash at any point in the
wake to the downflow at the airfoil decreases with in-
creasing Mach number and apex angle.


N-16329*

Aeronautical Research Council (Gt. Brit.)
THE AERODYNAMIC DERIVATIVES WITH RESPECT
TO SIDESLIP FOR A DELTA WING WITH SMALL
DIHEDRAL AT ZERO INCIDENCE AT SUPERSONIC
SPEEDS. A. Robinson and J. H. Hunter-Tod. 1952.
14p. diagrs. (ARC R M 2410; ARC 11,322. For-
merly College of Aeronautics, Cranfield. Rept. 12)

Expressions are derived for the sideslip derivatives
on the assumptions of the linearized theory of flow
for a delta wing with small dihedral flying at super-
sonic speeds. A discussion is included in the Appen-
dix on the relation between two methods that have
been evolved for the treatment of aerodynamic force
problems of the delta wing lying within its apex Mach
cone. When the leading edges are within the Mach
cone from the apex, the pressure distribution and the
rolling moment are independent of Mach number but
dependent on aspect ratio. When the leading edges
are outside the apex Mach cone, the nondimensional
rolling derivative is, in contrast to the other case,
dependent on Mach number and independent of aspect
ratio; the other derivatives and the pressure, how-
ever, are dependent on both variables.


N-16330"

Aeronautical Research Council (Gt. Brit.)
USE OF NEGATIVE CAMBER IN THE TRANSONIC
SPEED RANGE. W. F. Hilton. 1952. 5p. diagrs.
(ARC R & M 2460. Formerly ARC 10, 421; FM 1077;
S & C 2100)

Certain difficulties are experienced when attempting
free flight in the transonic speed range (0. 8 to 1. 2 of
the speed of sound). These difficulties fall into two
main classes: namely, (1) overcoming the somewhat
high air resistance by means of improvements in en-


NACA
RESEARCH ABSTRACTS N0.27


gine design, and (2) balancing the aircraft for stable
horizontal flight. The latter problem is considered
in this paper. Changes of trim are caused by sudden
loss of wing lift in the transonic range, which de-
creases the downwash over the tail, and possibly re-
sults in an uncontrollable nose-heavy dive. The use
of negative wing camber for minimizing these effects
is suggested. and the suggestion is found to be sup-
ported by wind-tunnel experiments.


N-16331*

Aeronautical Research Council (Gt. Brit.)
EXPERIMENTS ON THIN TURNING VANES. C.
Salter. 1952. 29p. diagrs., photo., 9 tabs. (ARC
R & M 2469. Formerly ARC 10, 039; TP 164)

The experiments recorded in this paper refer to a
variety of tests made on corner vanes (mainly thin
curved plates) in a tunnel 1 foot square, with the ob-
ject of indicating whether and to what extent they
would be suitable for deflecting the air stream around
the corners of return-flow type wind tunnels. It was
found that the air stream could be turned very satis-
fa( t'rilly by the use of thin vanes and that the loss of
head in a 90 corner was not as great as had been ex-
pected. Taking the pitch of the vanes to be the geo-
metrical pitch in a direction perpendicular to the axis
of the tunnel and the chord to be the distance between
the centers of radii of the leading and trailing edges,
the optimum ratio of pitch to chord for minimum loss
was shown to be about 0. 25. At that P/C ratio and
Reynolds number of 2 x 105, based on chord length
and air velocity of approach, the loss of total head in
the corner was found to be 0. 12 of the dynamic head,
while in the region outside the boundary layer the
minimum value was only 0. 06. To ensure stability
of the stream, a P/C ratio of 0. 2 is recommended
and it appears that, in order to maintain a thin bound-
ary layer, the actual number of vanes in any one cor-
ner should be not less than say 15 to 20. A clean en-
try of the streamlines into the vane unit can be estab-
lished by giving the leading-edge tangents an incidence
of about 5 relative to the upstream axis of the tunnel.
The minimum corner loss with 45 deflection was
found to be about 0.05 of the dynamic head with a
P/C ratio of roughly 0. 4. A set of "Collar" vanes
with splitters gave very satisfactory turning of the
stream with a loss of head in the corner of 0. 17 of the
dynamic head, at a Reynolds number of 150, 000, the
corresponding figure in the region outside the bound-
ary layer being 0.12. As the effect of the walls in
assisting the deflection of the stream is so large it is
suggestedthat the static-pressure distribution is as
good a criterion as any of the effectiveness of the cor-
ner. Uniformity of static pressure over both up-
stream and downstream cross-section to within a
short distance from the walls is a good test for the
correct shape and setting of the vanes.


N-16332*

Aeronautical Research Council (Gt. Brit.)
STRAIN GAUGE FLUTTER TESTS ON A 4-BLADE
PROPELLER WITH DURALUMIN BLADES.
J. Kettlewell and H. G. Ewing. 1952. 4p. diagrs.
(ARC R & M 2471; ARC 9876. Formerly RAE Tech.
Note SME 354)






NACA
RESEARCH ABSTRACTS N0.27


This note describes tests made on a fluttering duralu-
min propeller with strain gages fitted. Three dif-
ferent blade settings in the stalling region were in-
vestigated. The flutter encountered was purely tor-
sional; and the amplitude of strain recorded was at
no time serious.


N-16333*

Aeronautical Research Council (Gt. Brit.)
WIND-TUNNEL TESTS ON THE 30 PER CENT.
SYMMETRICAL GRIFFITH AEROFOIL WITH EJEC-
TION OF AIR AT THE SLOTS. N. Gregory, W. S.
Walker and W. G. Raymer. 1952. lip. diagrs.
(ARC R & M 2475. Formerly ARC 10, 097; FM 1018,
Perf. 251)

It has been shown by Preston (1946) that ejection of air
at the point of velocity discontinuity on a 16.2 percent
thick Griffith suction airfoil prevents separation, and
that if sufficient air is ejected, the drag is reduced.
The present tests were undertaken to apply this prin-
ciple to the 30 percent Griffith airfoil and to investi-
gate the effect on lift by pressure-plotting the airfoil.
Ejection of air was found to prevent separation, but
about 66 percent more air was required than with suc-
tion. Three times the suction quantity of air, when
ejected, reduced the drag to the low values associated
with suction. At R = 0. 96 million, the range of the
tests was 0-18 incidence and 0-14 flap angle. At
180 incidence and 14 flap angle, a CNF of 2. 5 was
obtained, giving approximately the same lift-curve
slope as with suction. Above this angle of incidence,
the pump capacity was not large enough for unsepa-
rated flow to be attained. With separation prevented,
the pitching moments were the same as with suction,
but the hinge moments were sensitive to small
changes of blowing quantity. At R = 2. 88 million, the
pump capacity was insufficient to prevent a partial
stall at 6 incidence as occurred with suction.
Curves of CNF, CQ, CM, CH, CD and velocity dis-
tribution when blowing are given, and comparisons
are made with corresponding curves obtained with
suction and with no suction. The same lift and pitch-
ing moments are obtained at any incidence with blow-
ing and with suction, but the suction quantities are
about 40 percent less than the blowing quantities.
The hinge moments are greatly different with blowing,
and increase of the normal force.


N-16334*

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF POINTED TIPS ON WING LOAD-
ING CALCULATIONS. V. M. Falkner. 1952. 4p.
diagr., tab. (ARC R & M 2483. Formerly ARC
10, 037; S & C 2063; Perf. 235)

The standard formulas for calculation of wing loading
by vortex lattice theory, which involve a sequence
starting with 4f( p2), are not strictly valid for
pointed wing tips. On the other hand, the use of a
sequence starting with 1 12 leads to overcorrection.
By considering the mean of these two solutions for a
delta wing, it is shown that the error introduced by
the use of the standard sequence is small and is on
the safe side as regards bending moments.


7


N-16335*

Aeronautical Research Council (Gt. Brit.)
FLIGHT MEASUREMENTS OF THE PRESSURE DIS-
TRIBUTION ON A TEMPEST WING UP TO A MACH
NUMBER OF 0. 8. K. Eyre. 1952. 15p. diagrs.
(ARC R & M 2489; ARC 10, 550. Formerly Hawker
Aircraft, Ltd. Design Rept. 1131)

This investigation was to determine the effect of
Mach number on the pressure distribution of the
Hawker airfoil section as fitted to Tempest and Fury
aircraft. The distributions show the typical super-
sonic flow conditions obtained over an airfoil of this
type. The airfoil does not show any unusual charac-
teristics and there are no large changes in pitching
moment and position of the aerodynamic center below
a Mach number of about 0. 75. In all cases the favor-
able pressure gradient in front of the shock wave is
fairly gradual, but changes rapidly at the shock wave.
There is, however, no sign of a breakdown in flow at
the trailing edge, where the distributions are similar
for all values of Mach number. The values obtained
for critical Mach number agree quite well with those
given in the Royal Aeronautical Society data sheets.
The variations of center-of-pressure position, pitch-
ing moment, and aerodynamic-center position are in
quite good agreement with German wind-tunnel tests
made on an approximately similar airfoil.


N-16336*

Aeronautical Research Council (Gt. Brit.)
24-FT. WIND-TUNNEL TESTS ON A "PADDLE-
BLADE" PROPELLER. A. B. Haines. 1952. 15p.
diagrs. (ARC R & M 2403; ARC 10, 308. Formerly
RAE Aero 2172)

The report contains the results of tests made in the
Royal Aircraft Establishment 24-foot tunnel on a de
Havilland propeller having conventional Clark Y sec-
tions and the paddle-blade type of plan form. The
report describes analyses of the results made both
by single-radius and full-strip theory methods. The
data provided will enable estimates to be made of the
performance of paddle-blade propellers under various
flight conditions and, in particular, will give a quan-
titative idea of the improvements to be expected for
high power loadings, that is, beyond the stall. The
principal general conclusions are: (1) Use of the
same lift/drag data in eight radius strip theory calcu-
lations for both paddle-blade and normal propellers
is justified. (2) The gains found in flight tests for
paddle-blade propellers must be principally ascribed
to the effects of NACA 16 series sections. (3) RAE
charts based on single-radius data may be pessimis-
tic in predicting the peak efficiencies of paddle pro-
pellers at high tip speed (by 1. 5 to 2 percent for
MT = 0. 85). (4) The mean stalling performance is
better than would be expected even on an activity fac-
tor basis. This enables other design modifications,
for example, thin roots, 16 series sections, etc. to
be made to improve the top speed performance with-
out causing serious losses at take-off or climb.
(5) Therefore, the general conclusion is that use of
the paddle type of plan form is fully justifiable even
though it does not appear to offer any direct shock-
stalling relief.







8



N-16337

Aeronautical Research Council (Git. Brit.)
TESTS ON THREE EQUILATERAL TRIANGULAR
PLATES IN THE COMPRESSED AIR TUNNEL.
R. Jones and C. J. W. Miles. 1952. 7p. diagrs.,
4 tabs. (ARC R M 2518. Formerly ARC 9973;
Perf. 222)

The report gives CL, CD, and Cm on three equi-
lateral triangular plates with sides 26, 36, and
47. 8 inches. The object of the tests was to estimate
the size of sweptback wings of that form, which can
be tested in the compressed air tunnel without the
application of excessive tunnel corrections. After
applying the usual tunnel corrections and making
small corrections due to dissymmetry in the models,
it was found that CL on the three models .,.., i:ed if
the results on the 36-inch and 47. 8-inch models were
multiplied by 1. 01 and 1. 05, respectively. Fair
.ILt*i -t liiCl was obtained on CD for the three models.
The mean position of the center of pressure is about
0.45 of the height of the tri.incl,., forward ot the trail-
ing edge.



N-16338*

Aeronautical Research Council (Gt. Brit.)
THE YAWING VIBR'ATIONS OF AN AIRCRAFT.
J. Morris and G. S. Green. 1952. llp. diagrs.
(ARC R '- M 2525; ARC 9482. Formerly RAE
SME 4027)

This report gives a theoretical method for calculating
the natural frequencies and modes of yawing vibration
of a complete aircraft. The basic feature of the
treatment is the replacement of the continuous mass
system by one consisting of a finite number of dis-
crete masses elastically interconnected. In the
course of the analysis, use is made of the deflection
coefficient artifice in the formation of the equations
of motion, and the escalator process in their mar-
shalling and numerical solution. The method has
been applied to a single-engined fighter aircraft, for
which the results of a resonance test were available.
These results appear to be some 40 percent in excess
of their calculated counterparts and no satisfactory
explanation occurs to the authors to account for this
incompatability.



N-16339*

Aeronautical Research Council (Gt. Brit.)
AIRCRAFT LANDINu GEAR: GROUND LOADS
WHEN SPINNING-UP THE WHEELS AT TOUCH-
DOWN. J. W. Blinkhorn. 1952. 13p. diagrs.,
2 tabs. (ARC R z M 2588; ARC 11, 627. Formerly
RAE Tech. Note Mech. Eng. 16)

The investigation covers all combinations of landing
speed, coefficient of friction between the tire and
the ground, and harshness of landing, for any type
of pneumatic tire and wheel unit. Particular atten-
tion has been given to landing speeds between 50 and
150 mph, coefficients of friction from 0 to 2. 0, and
landings giving vertical wheel accelerations of ig,


NACA
RESEARCH ABSTRACTS NO.27


2g, 3g, and 4g. It was found that for any landing, the
vertical reaction, at any wheel which has just finished
spinning up increases with increase in the moment of
inertia of the wheel and tire unit, and the landing
speed, and decreases with increase in the free tire
radius, the aircraft weight, the time to reach the
maximum vertical wheel reaction, and the coefficient
of friction between the tire and the ground. It should
be noted, of course, that there is a relation between
the moment of inertia, the free tire radius and the
aircraft weight in general the free tire radius and
the moment of inertia will increase with aircraft
weight. For any wheel and tire unit, it is shown that
there are various combinations of landing speed and
coefficient of friction which cause the wheel spinning
up to just cease at the same instant as the maximum
vertical wheel reaction is reached, and except for
very gentle l]ridirs, the maximum value of 1 re-
quired is usually much less than 1. 0.


N-163404


Aeronautical Research Council (Gt. Brit.)
THE SOLUTION BY LIF TINU-LINE THEORY OF
PROBLEMS INVOLVING DISCONTINUITIES.
V. M. Falkner. 1952. 23p. diagrs., 15 tabs.
(ARC R 9 M 2592. Formerly ARC 10, 922; S & C
2158; Perf. 366)

The report, which has been written as a preliminary
to a later account of similar work in lifting-plane the-
ory, describes how wing loading problems involving
discontinuities are solved by lifting-line theory. The
four discontinuities considered are (a) direction of
leading or trailing edge, (b) incidence, (c) two-
dimensional lift slope, and (d) chord. As the effects
of the first are of minor importance in lifting-line
theory, attention is mainly confined to the last three,
the solution being based on the use of a few terms of
a Fourier series in conjunction with special functions
tabulated elsewhere. The work is limited to straight
unyawed flight and includes lift, induced drag, and
pitching, rolling, and yawing moments, all with or
without deflected landing flaps and ailerons. The
method of formation of the equations, and the solu-
tions of a representative range of problems for a
hypothetical wing, including loading due to incidence,
symmetrical wing twist, uniform roll, and deflected
flaps and ailerons, are fully described. An indica-
tion is given of how induced drag and yawing-moment
calculations will later be simplified by the use of
special derived functions. Absolute values of wing
properties as given by lifting-line theory are usually
too high, but the specification of correction factors
for viscosity is beyond the scope of the report.


N-16341*

Aeronautical Research Council (Gt. Brit.)
CALCULATED LOADINGS DUE TO INCIDENCE OF
A NUMBER OF STRAIGHT AND SWEPT-BACK
WINGS. V. M. Falkner. WITH APPENDIX. Doris
Lehrian. 1952. 43p. diagrs., 43 tabs. (ARC
R S M 2596. Formerly ARC 11, 542; Perf. 449;
S & C 2223)







NACA
RESEARCH ABSTRACTS NO.27


In this report are collected lriether the calculated
aerodynamic loadings due to incidence of a number
of straight and sweptback wings. The calculations
follow in the main the routine described previously in
another report, but include additions concerned with
induced camber and induced drag. An additional in-
vestigation is made of the effect of the NACA camber
on the properties at zero lift of a rectangular wing of
aspect ratio 6. A table is given of loading functions
for use in auxiliary solutions when the wing plan has
a discontinuity of direction at an arbitrary position
along the span.



N-16342 *

Aeronautical Research Council (Gt. Brit.)
ON THE DESIGN OF AEROFOILS FOR WHICH THE
LIFT IS INDEPENDENT OF THE INCIDENCE. B.
Thwaites. 1952. 17p. diagrs. (ARC R & M 2612.
Formerly ARC 10, 294; FM 1057; Perf. 280)

It has been shown in R & M 2611 how lift may be ob-
tained on airfoils independently of the incidence. In
this paper mathematical processes are set out of de-
signing such airfoils to have specified velocity dis-
tributions at certain incidences and lift coefficients.
Approximate and exact methods are given, corre-
sponding to the methods employed in the design of
ordinary airfoils. Several shapes are worked out,
some of them being the product of ideas not given in
R & M 2611. A full discussion of the characteristics
of such airfoils is given.


N- 16343*

Aeronautical Research Council (Gt. Brit.)
SPEEDS AND NORMAL ACCELERATIONS OF BOE-
ING CLIPPER AIRCRAFT ON NORTH AND SOUTH
ATLANTIC ROUTES. D. T. Jones. 1952. 14p.
diagrs., 6 tabs. (ARC R & M 2633; ARC 11, 640.
Formerly RAE Structures 25)

This report presents results obtained from V-g re-
corders fitted to Boeing Clipper aircraft on the North
and South Atlantic routes between September 1944 and
May 1946. The records cover about 3, 300 flying
hours and show that the maximum speed recorded is
215 mph (I. A. S.) and the maximum upward and down-
ward accelerations are 2. 3g and -0. 3g, respectively.
The two main groups of records considered differ
from one another not only in respect of route but also
in seasonal conditions and in proportion of flights
made in wartime. Therefore, differences between
the results cannot be simply ascribed to differences
of route. It appears from the analysis that the maxi-
mum speed likely to be attained in a large flying time
is somewhat greater in one group (North Atlantic) than
in the other (South Atlantic) and that the maximum ac-
celerations on the other hand are likely to be less for
the former than the latter.


9


N-16353

Aeronautical Research Council (Gt. Brit.)
THE MEASUREMENT OF YAWING MOMENT OF AN
ELLIPTIC WING ON THE WHIRLING ARM. A. S.
Halliday. August 2, 1949. 8p. diagrs. (ARC
12, 497; S & C 2320)

The report describes tests on an elliptic wing on the
Whirling Arm to determine the yawing moment due to
a continuous rate of yaw. The results have been com-
pared with a theoretical estimation and the agreement
found to be satisfactory.


N-163551

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE CONDITION OF MATRICES. Olga
Todd. September 6, 1949. 3p. (ARC 12,561;
0.833)

The object of this note is to establish the following
theorem: let A be a real n x n nonsingular matrix
and A' its transpose. Then AA' is more "ill-
conditioned" than A. This result confirms an opin-
ion expressed by Dr. L. Fox which he based on his
practical experience. The term "condition of a ma-
trix" has been used in vague senses. The most com-
mon measure of a matrix has been the size of its de-
terminant: ill-conditioned matrices being those with
small determinants. With this interpretation, the
theorem is clearly correct. More adequate meas-
ures of the condition of a matrix have been proposed
recently by John von Neumann and H. H. Goldstine
and A. M. Turing.


N-16356*

Aeronautical Research Council (Gt. Brit.)
A NOTE ON THE DESIGN OF A HIGH-SPEED WIND
TUNNEL DRIVEN BY STEAM. T. R. Cave-Browne-
Cave. September 15, 1949. 5p. diagrs. (ARC
12,577; TP 292)

A steam-driven supersonic wind tunnel has been de-
signed for, and is being built by, the University Col-
lege in Southampton. It has a 6 inch by 2-1/2 inch
working section and will operate at a Mach number of
1. 4 for a period of 3 minutes. Since it takes about 5
minutes for the boiler to return to pressure, a 3 min-
ute test could be made every 8 minutes with this set-
up. Because the boiler is already available at the
school, this design permits a substantial saving in
time and money. Another advantage will be the com-
parative lack of noise as compared with a drive using
a compressor.



N-16459*

Royal Aircraft Establishment (Gt. Brit.)
THE RHEOLOuY OF LUBRICANT FILMS. PART 2.
E. Bielak and E. W. J. Mardles. February 1952.
27p. diagrs. (RAE Chem. 480)




UNIVERSITY OF FLORIDA


3 1262 09079 7498
NACA
10 RESEARCH ABSTRACTS N0.27


This report presents further results on the rheolo-
gical properties of thin lubricant films (including
greases) moving between surfaces, with special re-
card to the influence of the material of the surfaces
on the rate of flow. The limitation of Amontons'
law to pressures above 100 g/sq cm for lubricants is
described in relation to the phenomenon of the "wring-
inf" together or "striction" of surfaces at light pres-
sures, especially with indifferent lubricants, when
film rupture occurs early during the thinning of the
film. The latent period, observed and first des-
cribed by Hardy, during which friction changes occur
has been measured for several lubricants at low pres-
sures. Explanations of hydraulic "striction" and the
"latent period" are suggested.


MISCELLANEOUS


NACA TN 2310

Errata No. I on "GENERALIZATION OF BOUNDARY-
LAYER MOMENTUM-INTEGRAL EQUATIONS TO
THREE-DIMENSIONAL FLOWS INCLUDING THOSE
OF ROTATINu SYSTEM". Artur Mager. March
1951.


NA"A-T.2naPleV 8-15-52 -4000




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