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fNational Advisory Committee for Aeronautics
i:a: Research Abstracts
iO.S FEBRUARY20, 1953
SACIA Rept. 1069
N SOLUTIONN OF THE NONLINN-
'L EQUATION FOR TRANSONIC ;O PAST A
tAVE-SHAPED WALL. Carl Kapla j iA. ,
:.lp. diagrs., tab. (NACA Rept. 1069. For -
h e simplified nonlinear differential equation for
trao~lc flow past a wavy wall is solved by the
method of integration iti series. The solution has
ui:been carried to the point where'the question of the
tftence dr rdonexistence of a mixed potential flow
icai be answered by the behavior of a single power
~ selies in the transonic similarity parameter. The
a, culation of the coefficients of this dominant power
alries has been reduced to a routine computing
protdinm by means of recursion formulas resulting
rIlom the solution of the differential equation and the
i~.1andary condition at the surface of the wavy wall.
VNAA Rept. 1071
: THEORETICAL SYMMETRIC SPAN LOADING DUE
TO FLAP DEFLECTION FOR WINGS OF ARBITRARY
PLAN FORM AT SUBSONIC SPEEDS. John DeYoung.
952. ii, 41p. diagrs., tabs. (NACA Rept. 1071.
Formerly NACA TN 2278)
p .' l cedure based upon a simplified lifting-surface
:; theory that includes effects of compressibility and
% !j:.::panwise variation of section lift-curve slope is pre-
Scented in such a manner that the spanwise loading due
Sto-flap deflection can be simply found for wings hav-
P. :j symmetric plan forms with constant spanwise
i'.~iaweep angle of the quarter-chord line. Aerodynamic
characteristics due to flap deflection are considered
k. .la sidlor straight-tapered wings, values of certain of
': these characteristics are presented in charts for a
Z;; range of swept plan forms. Further use of the meth-
od gives downwash in the vertical center of the wake
!: of the wing.
4' ; .ACA Rept. 1083
': AXJSYMMETRIC SUPERSONIC FLOW IN ROTATING
; IMPELLERS. Arthur W. Goldstein. 1952. ii,
., 14p, diagrs. (NACA Rept. 1083. Formerly
General equations are developed for isentropic,
f,: .rictionless, scisymmetric cbmpressible flow in
S rotating impellers with blade forces eliminated in
: r"favor of the blade-sutface function. The character-
i., tic equations for supersonic flow are developed
88LE ON LOAN ONLY.
IR UESTS FOR DOCUMENTS TO NACA, 1794 F ST., NW.,
)RT Ti LE AND AUTHOR.
t'i ".- a c;'.. ." .!*..
and a computing technique is utilized to find the ef-
fect of variations of design parameters on internal
flow and work-input distribution.
NACA TN 2881
AERODYNAMIC CHARACTERISTICS OF A TWO-
BLADE NACA 10-(3)(062)-045 PROPELLER AND OF
A TWO-BLADE NACA 10-(3)(08)-045 PROPELLER.
William Solomon. January 1953. 53p. diagrs.,
photo., tab. (NACA TN 2881. Formerly RM L8E26)
Characteristics are given for the two-blade NACA
10-(3)(062)-045 propeller and for the two-blade
NACA 10-(3)(08)-045 propeller over a range of ad-
vance ratio from 0.5 to 3.8, through a blade-angle
range from 200 to 550 measured at the 0.75 radius.
Maximum efficiencies of the order of 91.5 to 92 per-
cent were obtained for the propellers. The propeller
with the thinner airfoil sections over the outboard
portion of the blades, the NACA 10-(3)(062)-045 pro-
peller, had lower losses at high tip speeds, the
difference amounting to about 5 percent at a helical
tip Mach number of 1.10.
NACA TN 2884
CALCULATION AND MEASUREMENT OF NORMAL
MODES OF VIBRATION OF AN ALUMINUM-ALLOY
BOX BEAM WITH AND WITHOUT LARGE DIS-
CONTINUITIES. Frank C. Smith and Darnley M.
Howard, National Bureau of Standards. January
1953. 40p. diagrs., photo., 8 tabs. (NACA TN 2884)
The lowest normal modes of vibration of three alumi-
num alloy box beams were calculated using a matrix
iteration method. For the calculations the actual
structures were idealized to a system of mass points
interconnected by massless springs. The lowest
normal modes ol these beams were measured experi-
mentally and compared with those calculated. This
comparison indicates that the mode shapes and
natural frequencies for structures of this type may be
adequately calculated using this method. The experi-
mental measurements were limited at the higher
frequencies by local vibrations of small elements of
NACA TN 2885
SOME EXACT SOLUTIONS OF TWO-DIMENSIONAL
FLOWS OF COMPRESSIBLE FLUID WITH HODO-
GRAPH METHOD. Chieh-Chien Chang and Vivian
O'Brien, Johns Hopkins University. February 1953.
63p. diagrs., 4 tabs. (NACA TN 2885)
WASHINGTON 95. D. C, CITING CODE NUMBER ABOVE EACH TITLE;
A suggestion is given for classifying compressible
potential flows according to location and number of
singularities in the subsonic region of the hodograph
plane, which seems to offer a convenient, criterion
for systematic investigation of these flows with
Chaplygin's original method. The object of the pa-
per is to present and analyze a few useful solutions
of compressible potential flow with the exact gas
law. These solutions include flows about convex
corners and belong to the same class as that of
Ringleb. Also, the exact solution of compressible
flow through a particular contracting channel is
NACA TN 2886
AN ANALYSIS OF STATICALLY INDETERMINATE
TRUSSES HAVING MEMBERS STRESSED BEYOND
THE PROPORTIONAL LIMIT. Thomas W. Wilder,
II. February 1953. 13p. diagrs., 4 tabs. (NACA
A procedure for analyzing statically indeterminate
trusses in the plastic stress range is presented which
is applicable to trusses having any number of redun-
dant members. By using the Ramberg-Osgood ana-
lytical representation of the stress-strain curve, the
analysis of the truss is reduced to the solution of a
set of simultaneous equations. A numerical example
is presented to illustrate the procedure.
NACA TN 2888
PERFORMANCE CHARACTERISTICS OF PLANE-
WALL TWO-DIMENSIONAL DIFFUSERS. Elliott G.
Reid, Stanford University. February 1953. 1, 80p.
diagrs., photos., 3 tabs. (NACA TN 2888)
Performance characteristics were determined for
plane-wall, two-dimensional diffusers which were so
proportioned as to insure reasonable approximation
of two-dimensional flow. The diffusers had identical
entrance cross sections and discharged directly into
a large plenum chamber; the test program included
wide variations of divergence angle and length. A
dynamic pressure of 60 pounds per square loot was
maintained at the diffuser entrance and the boundary
layer there was thin and fully turbulent. A few tests
were made with asymmetric diffusers. Others
showed the effects of addition of a short exit duct of
uniform section and of installation of a thin, central,
NACA TN 2891
FACTORS AFFECTING LAMINAR BOUNDARY LAY-
ER MEASUREMENTS IN A SUPERSONIC STREAM
Robert E. Blue and George M. Low. Appendix B:
REDUCTION OF DATA. Jack M Lande. February
1953. 49p. diagrs. (NACA TN 2891)
The observed discrepancy at supersonic speeds be-
tween theoretical and apparent experimental average
flat plate Iriction-drag coefficients calculated from
boundary layer total-pressure surveys was investi-
gated. Effects of the total-pressure probe, heat
transfer through the leading-edge region, leading-
RESEARCH ABSTRACTS NOiI :' ..
edge geometry and strength of the leadlmg-edfle,
possible early transition to turbulent flow drb "b
of turbulence, and the slight streamwil tift
gradient inherent in flat-plate flow were i li
the investigation. Only one of these fact, 'tfhe:.
feet of the total-pressure probe, was foudti..e -...
significant. Total-pressure probes of different tip
heights, when placed in laminar boundary la yeiAe-
veloping under identical conditions, measure dz er-
ent values of friction-drag coefficient. EG"rapolk-,
tion of these measurements indicates that sa .*u of
vanishing tip height would measure the theoreticall;y
predicated values of average flat plateffriction-d ig
NACA TN 2894
CALCULATIONS OF UPWASH IN THE REGION
ABOVE OR BELOW THE WING-CHORD PLANESOF. .
SWEPT-BACK WING-FUSELAGE-NACEL.LE. COM-
BINATIONS. Vernon L Rogallo and.John-L. '
McCloud, m. February 1953. 15p. diagra-, photo".'
(NACA TN 2894)
A procedure has been developed for predicting theW..
upwash components of the upflow angles in the reglt.i.
above or below the wing-chord planes of swept-back .ifl
wing-fuselage-nacelle combinations. Compariso.ns
are made of the predicted and measured upflow an- :.
gles for six semispan models with 4Q swept-back '
NACA TN 2896
SURVEY OF PORTIONS OF THE IRON-NICKEL- .
MOLYBDENUM AND COBALT-IRON-MOLYBDEN IM
TERNARY SYSTEMS AT 12000 C. Dilip K. Da an d':
Paul A. Beck, University of Notre Dame. :'
February 1953. 56p. diagrs., photos., I& tabs. ::
(NACA TN 2896)
The 12000 C isothermal sections of the iron-nickel
molybdenum and the cobalt-iron-molybtenuni teparay
systems were surveyed. The phases occurringn.
these systems were identified by means ot X-ray-'
diffraction and by etching methods, and the phase'
boundaries at 12000 C were determined microlIeopi-''
cally, using the disappearing phase method with' '
quenched specimens. Both systems contain long :
solid-solution fields of the mu phase. Other inter-- .
mediate phases occurring in the iron-nickel- ::
molybdenum system are the P phase and the delta
phase. Both phase diagrams have extensive face-
centered cubic solid-solution fields and some body-
centered cubic solid solutions.
NACA RM E52L09-
FORCED-CONVECTION HEAT-TRANSFER CHAR- .
ACTERISTICS OF MOLTEN SODIUM HYDROXIDE ...
Milton D. Grele and Louis Gedeon. February. 1953,.,
27p. diagrs., photo., 2 tabs. (NACA RM E52LO.B).- ..
The forced-convection heat-transfer characteristics ::
of sodium hydroxide were experimentally investi-
gated. The heat-transfer data for heating fall slghtl '.
RESEARCH ABSTRACTS NO.38
ly above the McAdams correlation line, and the heat-
transfer data for cooling are fairly well represented
by the McAdams correlation line.
I qACA TM 1344
ION THE THEORY OF THE TURBULENT BOUNDARY
.AYER. (Uber die Theorie der turbulenten
i' renzachichten). J. Rotta. February 1953. 50p.
fagrs. (NACA TM 1344. Trans. from Max-
'latick-Institut fTr Stromungslorschung, G6ttingen.
i1tteilungen 1, 1950)
turbulent energy, dissipation, and momentum rela-
Stionsare discussed. A procedure is given for com-
puttation of turbulent skin friction in boundary-layer
.flow -ith pressure gradients. The boundary layer is
divided into a region near the wall where viscosity
and surface roughness are important, an outer re-
gion which is dependent on friction coefficient and
pressure gradient, and an intermediate zone between
these two which is unaffected by wail roughness, vis-
"cosily, and the outer flow. Analytical confirmation
is obtained for the empirical fact that turbulent
boundary layers are able to overcome a greater
pressure rise than laminar ones.
NACA TM 1349
ON A CLASS OF EXACT SOLUTIONS OF THE
EQUATIONS OF MOTION OF A VISCOUS FLUID.
(Ob odnom klasse tochnykh reshenii uravnenii
dvizheniya vyazkoi zhidkosti). V. I. Yatseyev
February 1953. 7p. (NACA TM 1349. Trans.
from Zhurnal Eksperimental 'noi i Teoretisheskoi
Fiziki, v.20, no. 11, 1950, p. 1031-1034).
Thie general solution is obtained of the equations of
motion of a viscous fluid in which the velocity field
is.inversely proportional to the distance from a
certain point. -Some particular cases of such mo-
tion are investigated.
NACA TM 1358
CALCULATION OF THE SHAPE OF A TWO-
DIMENSIONAL SUPERSONIC NOZZLE IN CLOSED
FORM. (Sul Calcolo in Termiru Finiti dell'Effusore
di una Galleria Bidimensionale Supersonica). Dante
Cunsolo. January 1953. 29p. diagrs. (NACA
TM 1358. Trans. from Aerotecnica, v.31, no.4,
August 15, 1951, p. 225-230).
The idea is advanced of making a supersonic nozzle
by producing one, two, or three successive turns of
the whole flow; with the result that the wall contour
can be calculated exactly by means of the Prandtl-
Meyer "Lost Solution. '
Aeronautical Research Council IGt. Brit.)
NOTE ON PROFILE DRAG CALCULATIONS FOR
LOW-DRAG WINGS WITH CUSPED TRAILING
EDGES. R. C. Lock 1952. 10p. diagrs., tab.
(ARC R & M 2419; ARC 9772. Formerly RAE
In R. M. 1833 calculations of profile drag were
made based on wing sections of conventional design,
and were later extended in an Addendum to "low-
drag" wing sections with convex trailing edges.
Further calculations were required for low-drag
sections of more recent design with cusped trailing
edges. This report presents calculations made on
sections of the NACA 65-family of thickness 0. 12c
and 0 23c with maximum thickness at 0. 4c from the
leading edge, over a range of Reynolds number and
position of the transition points.
Aeronautical Research Council lGt Brit.)
MODEL TESTS WITH FLOW ON THE GLOSTER
F.9 40 WITH H. 1 NACELLES (METEOR H). J. S.
Thompson. C. M. Fougere and E. G. Barnes. 1952.
17p. diagrs., 14 tabs. (ARC R M 2517; ARC 6901.
Formerly RAE Aero 1821)
For an earlier estimate by Thompson and Barnes of
the longitudinal stability of the F. 9 40 in flight, it
was necessary to extrapolate for the higher values
of CL, because the maximum jet flow from model
tests then available was that appropriate to a CL of
only 0. 2, ground level. The purpose of the tests
described in this note was to extend the previous
model tests, using a considerably larger flow, to
enable more precise estimates to be made. Meas-
urements were made of lift, drag, pitching and yaw-
ing moments for jet flows up to an exit v V of 4. 5
and at various tunnel speeds.
Aeronautical Research Council (Gt. Brit.)
THE DESIGN AND INSTALLATION OF SMALL COMA
PRESSED AIR TURBINES FOR TESTING POWERED
DYNAMIC MODELS IN THE ROYAL AIRCRAFT
ESTABLISHMENT SEAPLANE TANK. D. I. T. P.
Llewelyn-Davies, W. D, Tye and D. C. MacPhail.
1952. 19p. diagrs.. photos., tab. (ARC
R & M 2620; ARC 10.812 Formerly RAE
This report describes the development of small light-
weight air turbines for powering dynamic models in
the R. A. E. seaplane tank. The units have proved to.
be rugged and reliable and power., weight ratios of 0.4
lb bhp have been achieved. The installation of the
turbines in dynamic models and the provision of their
air supply are also discussed.
Aeronautical Research Council (Gt. Brit.)
ON THE SOLUTION OF LINEAR SIMULTANEOUS
DIFFERENTIAL EQUATIONS WITH CONSTANT
COEFFICIENTS BY A PROCESS OF ISOLATION.
J. Morris. 1952. 7p. (ARC R & M 2623;
ARC 11,420. Formerly RAE SME 4036)
In this report, a process is given for the solution of
linear differential equations with constant coeffi-
cients. The operative artifice is closely akin to
Routh's method of isolation by means of which the
constants of integration are found separately for each
root of the characteristic equations.
Aeronautical Research Council (Gt. Brit.)
THE DYNAMIC LANDING LOADS OF FLYING
BOATS WITH SPECIAL REFERENCE TO MEASURE-
MENTS MADE ON SUNDERLAND TX. 293. Anne
Burns and A. J. Fairclough. 1952. 38p. diagrs.,
photos., 2 tabs. (ARC R M 2629; ARC 11,344.
Formerly RAE Structures 17)
An account is given of a full-scale investigation into
the stresses occurring in the wing members of a
Sunderland flying boat during landing impacts. It is
found that the main dynamic effect is caused by the
wing oscillating in its fundamental mode. These
dynamic loads have a spanwise distribution similar
to the normal lift load and, if the level flight lift load
is taken as unity, a magnitude (in the most severe
impact recorded) of 1.4 upwards and 1.5 downwards.
Generalizing this result, one concludes that whereas
down loads in landing may be a deciding factor in
design the up loads are amply covered by existing
requirements. Comparison of calculated and exper-
mental loads found in these tests indicates that sat-
isfactory agreement can be attained by using recent-
ly introduced modifications of standard dynamical
methods. Although the investigation is primarily a
structural one some interesting results on general
water load phenomena are obtained.
Aeronautical Research Council (Gt. Brit.
LOAD DIFFUSION AT AN INTERSPAR OPENING:
THEORETICAL METHODS OF ANALYSIS COM-
PARED WITH STRAIN MEASUREMENTS ON A
LARGE WING. D. C. Allen. 1952. 26p diagrs.,
6 tabs. (ARC R r M 2664; ARC 11,731. Formerly
RAE Structures 30)
The diffusion of load from spar flanges into skin and
stringers near an opening was investigated experi-
.mentally in a large wing structure undergoing
strength tests. A comparison of measured strains
with those given by theoretical methods shows that in
general the flange loads are represented with rea-
sonable accuracy. Any theory, however, in which
the chordwise rib at the edge of the opening is ig-
nored gives shear stresses much greater than those
measured. Allowance for the bending stiffness of
this rib produces values of shear stress comparable
with those obtained experimentally.
RESEARCH ABSTRACTS NO.8
N-20668* 1 ,: :
Aeronautical Research Council (Gt. Brit.) ".
LANDING GEAR WITH TWIN TANDEM WHEq]L::. .
UNITS: CORNERING CHARACTERISTICS AS DE-: A
TERMINED BY MODEL TESTS. J. W..Blii horn...
1952. 7p. diagrs., photos., tab. (ARC R&. M2668;
ARC 11,850. Formerly RAE Tech. Note Meeh. Eng.
For twin tandem units, the wheel loading conditions
which arise when aircraft are turned on the ground : i'
may be critical for the landing gear. To estimate .:,
the magnitude of these loads, cornering tests wer ...
made on a small-scale model of the main under-.
carriage unit proposed for the Brabazon 1, MK. II. .
These tests showed that for zero turning radius, that
is, turning about the central vertical axis of the
model undercarriage, the wheel side loads were al-
most equal to the vertical load multiplied by the '
coefficient of sliding friction between the tires and
the ground. The side loads rapidly decreased as the .
turning radius increased, and with the turning radius" "'
equal to three times the wheel base, the wheel side
loads were only about half of those at zero turning
radius. The severity of the design loads for turning
on the ground will therefore be considerably reduced .:
if it can be ensured that the center of thd minimum
turning circle of the aircraft is a short distahe ".
outboard of either main undercarriage unit.
Aeronautical Research Council (Gt. Brit.)
CONICAL FLOW AS A RESULT OF SHOCK AND ',
BOUNDARY-LAYER INTERACTION ON A PROBE.
J. Lukasiewicz. 1952. 16p. diagrs., photos..
(ARC R & M 2669; ARC 12, 023. Formerly RAE ..
Tech. Note Aero 1968; SD 85) .
The formation of a conical shock and a conical region..
of flow separation originating from the tip of a thin
traversing tube was observed in a supersonic tunnel
as a result of interaction of a strong shock with the.
boundary layer on the tube surface. The angles of .
the conical shock and separation surfaces and the
static pressure in the separation region are in good .
agreement with the theoretical conical flow solutions.. .'
The extent of the conical flow illustrated should act .
as a warning against the use of static pressurs tube'i;.:
for measuring pressures in the regions of strong .
Aeronautical Research Council (Gt. Brit.)
AN ELECTRIC TANK FOR THE DETERMINATION
OF THEORETICAL VELOCITY DISTRIBUTION
T. J. Hargest. 1952. 9p. diagrs., photos. .A kC .:;
R& M 2699; ARC 12,448. Formerly NOTE Miemo.:
An analogy due to Relf has been applied to the design :...
of apparatus for quickly determining the theoretici- ::.
velocity distributions around an airfoil in cascade.'
The accuracy of the apparatus was tested by deteri"'^.
mining the velocity distribution around a cylinder:' t'
'," .. ..:
RESEARCH ABSTRACTS NO.38
An accuracy of within 1 percent of the approach
velocity was obtained for this case. The apparatus
.: has since been applied to determine the theoretical
: velocity distribution around various airfoils in cas-
r cades; an example is given of the pressure distribu-
tion around an airfoil at zero incidence. An applica-
tion to determine the theoretical velocity distribution
around the central airfoil of a nozzle cascade where
the effect of the ducting side walls is included is also
Aeronautical Researoh Council (Gt. Brit.)
VELOCITY DISTRIBUTION ON STRAIGHT AND
SWEPT-BACK WINGS OF SMALL THICKNESS AND
INFINITE ASPECT RATIO AT ZERO INCIDENCE.
S. Newmark. 1952. 40p. diagrs. (ARC R M 2713;
ARC 10, 907. Formerly RAE Aero 2200)
A solution by H. Ludwieg, giving the velocity distri-
bution in the central section of a thin sweptback wing
of infinite aspect ratio with a biconvex profile at zero
incidence, has been found erroneous. In connection
with this problem, the approximate method of
sources and sinks for determining velocity distribu-
tion on straight and sweptback wings is critically
examined, its limitations established, and proper
ways of its application to three-dimensional problems
indicated. A correct solution of Ludwieg's problem
is found, and generalized to give the velocity distri-
bution over the entire wing. The method is further
extended to cover a wide class of thin symmetrical
wing profiles, those with rounded leading edge being,
however, often intractable by this particular method.
The ultimate purpose of the investigation is to pro-
vide a reliable basis for determining the critical
Mach number for sweptback wings. Further work
is needed to embrace wings of finite aspect ratio and
tapered wings, in particular, delta wings The
method seems adequate to deal with these more com-
Aeronautical Research Council (Gt. Brit
THE NUMERICAL SOLUTION OF TWO-
DIMENSIONAL FLUID MOTION [N THE NEIGH-
BOURHOOD OF STAGNATION POINTS AND SHARP
CORNERS. L. C. Woods. 1952. 15p. diagrs.
(ARC R & M 2726. Formerly ARC 12, 887, FM 1407)
Methods are given in this paper of dealing with sin-
gularities of functions satisfying certain two-
dimensional partial differential equations. For a
numerical solution, the differential equations are
replaced by difference equations on a square mesh.
Log (1/q) where q is the velocity, becomes infinite at
stagnation points, sharp corners, sinks, etc., while
the conjugate function a (flow direction) becomes
multivalued. The method consists in finding a
series expansion for the function (log 1. q or 0) in the
neighborhood of the singularity. This expansion is
then used to find relationships between the function
values at points of the mesh adjacent to the singulari-
ty. A method of working directly in the transformed
flow plane tin which the airfoil is a slit), and thus
avoiding irregular squares on the boundary, is also
given. The method is developed for incompressible
flow, but an approximation suitable for compressible
flow is given.
DECLASSIFIED NACA REPORTS
NACA RM A8E17
THE ASYMMETRIC ADJUSTABLE SUPERSONIC
NOZZLE FOR WIND-T UNNEL APPLICATION.
H. Julian Allen. July 23. 1948. 42p. diagrs,
photos., 2 tabs. (NACA RM A8E171 (Declassified
from Restricted, 10 30 52)
An asymmetric adjustable nozzle ior supersonic wind
tunnel application which permits continuous adjust-
ment of the test-section Mach number is described.
The characteristics of this nozzle are compared with
the more conventional supersonic tunnel nozzles.
NACA RM L8E06
AERODYNAMIC CHARACTERISTICS AT HIGH
SPEEDS OF RELATED FULL-SCALE PROPELLERS
HAVING DIFFERENT BLADE-SECTION CAMBERS.
Julian D. Maynard and Leland B. Salters, Jr.
August 31, 1948. 54p. diagrs., tab., photo (NACA
RM L8E06) (Declassified irom Restricted,
11 26 52)
Comparisons are made of results obtained in wind-
tunnel tests of related full-scale propellers over a
range of blade angles from 200 to 550 at airspeeds
up to 500 miles per hour to evaluate the combined
effects of blade-section camber and compressibility
on propeller aerodynamic characteristics
NACA RM L8E07
AERODYNAMIC CHARACTERISTICS AT HIGH
SPEEDS OF FULL-SCALE PROPELLERS HAVING
CLARK Y BLADE SECTIONS. Peter J. Johnson.
October 26, 1948. 60p. diagrs photos., tab
(NACA RM L8E07) (Declassified from Restricted,
II 26 52)
Results obtained in wind-tunnel tests of two full-
scale propellers over a range oi blade angles from
200 to 550 and at airspeeds varying from 60 to 485
miles per hour are presented
NACA RM L8E24
AERODYNAMIC CHARACTERISTICS OF A TWO-
BLADE NACA 10-(3)(08)-03R PROPELLER.
Albert J. Evans and Leland B. Salters, Jr.
September 2, 1948. 29p. diagrs., tab. (NACA
RM L8E24) (Declassified from Restricted,
Contains results of wind-tunnel tests on a full-scale
NACA 10-(3)(08)-03R two-blade propeller The
w war ar
UNIVERSITY OF FLORIDA
3 1262 08153 270
tests were part of a program to determine the effects
of blade-shank design on propeller aerodynamic
characteristics. A maximum efficiency of 91.5 per-
cent was attained at a rotational speed of 1600
revolutions per minute at a 300 blade angle. Peak
efficiency at a blade angle of 450 was decreased 32
percent by increasing the hclical-tip Mach number
from 0.80 to 1.20.
NACA RM L8H16
AERODYNAMIC CHARACTERISTICS OF A THREE-
BLADE PROPELLER HAVING NACA 10-(3)(08)-03
BLADES. Robert E. Davidson. October 29, 1948.
29p. diagrs., tab. (NACA RM L8H16) (Declassified
from Confidential, 11 26, 52)
Contains results of wind-tunnel tests of a 10-foot-
diameter, three-blade propeller at stream Mach
numbers from 0.12 to 0.64. The tests were part of
a program to determine the effects of blade number
on propeller aerodynamic characteristics. The pro-
peller blades are of designation NACA 10-(3)(08)-03.
A maximum efficiency of 92 percent was attained at
a rotational speed of 1350 rpm at a 400 blade angle.
Peak efficiency at the design blade angle of 450 was
decreased 4 percent by increasing the helical tip
Mach number from 0.80 to 0.96.
NACA RM L9G20
TWO-DIMENSIONAL WIND-TUNNEL INVESTIGA-
TION OF A 6-PERCENT-THICK SYMMETRICAL
CIRCULAR-ARC AIRFOIL SECTION WITH
LEADING-EDGE AND TRAILING-EDGE HIGH-
LIFT DEVICES DEFLECTED IN COMBINATION.
Robert J. Nuber and Gail A. Cheesman.
September 6, 1949. 29p. diagrs., photo.. 3 tabs.
(NACA RM L9G20) (Declassified from Restricted,
An investigation ol a 6-percent-thick symmetrical
circular-arc airfoil with leading-edge and traillng-
edge high-lift devices was made to determine the
effectiveness of these devices in increasing the
maximum section lift coefficient of the airfoil when
deflected in combination. The results indicated that,
with the plain trailing-edge flap deflected 600, maxi-
mum section lift coefficients of 2.02 and 1.95 can be
obtained by deflecting a 15-percent-chord leading-
edge slat or a 15-percent-chord drooped-nose flap,
respectively. Variations in Reynolds number for
either the slat or drooped-nose-flap configurations
or moving the drooped-nose-flap hinge from the
lower surface to the upper surface had essentially
no effect on the lift characteristics.
NACA ..* '
RESEARCH ABSTRACTS -NO.="-f;..l
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