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THE REPORT TITLE AND AUTHOR.
Measurements are presented at Mach numbers from
about 1.3 to 1. 5 of reflection characteristics and the
relative upstream anfluence of shock waves rmpinging
on a flat surface with both laminar and turbulent
boundary layers. The difference between impulse
and step waves Is discussed and their Interaction
with the boundary layer as compared. General
considerations on the experimental pou
shock waves from wedges and cones at *
reflection of shock waves from supep 's hear
layers are also presented.
A METHOD FOR STABILIZING SHOCK WAVES IN
CHANNEL FLOW BY MEANS OF A SURGP-- --
CHAMBER. Stanford E. Nelce, Cornet ,University
June 1953. 46p. photos., diagrs. (NACA T'N 2694)
In order to stabdlize normal-shock waves In channel
flow against the effect of disturbances originating
downstream, a method based on mass removal from
the channel by means of a surge chamber was
developed and experimentally tested in an intermat-
tent blowdow~n-type wind tunnel at Cornell Uruversity.
A theoretical analysis was applied to a channel
shape similar to that used In a typical supersonic
wind tunnel and Indicated that the mass-removal
technique was effective In damping the motion of the
normal shock caused by a strong pulse originating
downstream. Experlmental results were in direct
agreement with the theoretical analysis. Further
ar~rmnl intndicaent that th ms-erno tor
motion of the normal shock caused by continuous
small, random, downstream disturbances.
NACA TN 2933
BEHAVIOR OF MATERIALS UNDER CONDITIONS
OF THERMAL STRESS. S. S. Manson. July 1953.
105p. diagrs., photos., 6 tabs. (NACA TN 2933)
National Advisory Committee For Aeronautics
JULY 10 1953
CURRENT NACA REPORTS
NACA Rept. 1086
ANALYSIS OF THE EFFECTS OF WING INTERFER-
ENCE ON THE TAIL CONTRIBUTIONS TO THE
ROLLING DERIVATIVES. William H. Michael, Jr.
1952. ii, 12p. diagrs. (NACA Rept. 1086.
Formerly TN 2332)
An analysis of the effects of wing interference on the
tail contributIons to the rolling stability derivatives
of complete airplane configurations is made by cal-
culating the angularity of the air stream at the verti-
cal tail due to rolling and determining the resulting
forces and moments. Some sidewash results are
presented for a limited range of wing plan forms and
vertical-tall sizes. Equations for using the sidewash
results to determine the lail contributions to the
rolling derivatives are given and some comparisons
with experimental data are made.
NACA Rept. 1092
FLIGHT INVESTIGATION OF THE EFFECT OF
CONTROL CENTERING SPRINGS ON THE
APPARENT SPIRAL STABILITY OF A PERSONAL-
OWNER AIRPLANE. John P. Campbell, Paul A.
Hunter, Donald E. Hewes and James B. Whitten.
1952. ii, 17p. diagrs., photo., 2 tabs. (NACA
Rept. 1092. Formerly TN 2413)
A flight investigation has been conducted on a typical
high-wing personal-owner airplane to determine the
:"":'.""h 01nro centee s rinings on a paent spth l
the aileron and rudder control systems to provide
both a positive centering action and a means of trim-
ming the airplane.
NACA Rept. 1100
ON REFLECTION OF SHOCK WAVES FROM
BOUNDARY LAYERS. H. W. Liepmann, A. Roshko
and S. Dhawan, California Institute of Technology.
1952. ii, 29p. diagrs., photos Lab. INACA
Rept. 1100. Formerly TN 2334)
RESEARCH ABSTRACTS NO.45
differetrial equation, suitable for use with a high-
speed digital computer, is described, and typical
co~mpue results are given. The existence of a
finite lifetime, although not evident from the
differential equation, is argued intuitively and con-
firmed by the numerical computations.
NACA TN 2957
SURVEYS OF THE FLOW FIELDS AT THE
PROPELLER PLANS OF SIX 400 SWEPTBACK
'WING- FUSELAGE-NACELLE COMBINATIONS.
Vernon L. Rogallo and John L. McCloud III. June
1953. 57p. diagrs., photos. (NACA TN 2957)
The flow fields at the propeller planes of six 400
swreptback, semispan wrng-[uselage-nacelle combi-
nations have been surveyed. The significance of the
flow-field parameters an terms of propeller-
oscillating aerodynamic loads is indicated. Compar-
isons of measured and predicted upflow angles are
made for all six models.
NACA TN 2958
REACTION PROCESSES LEADING TO SPONTANE-
OUS IGINITIOIN OF HYDROCARBONS. Charles E.
Frank and Angus U. Blackham, University of
Cincinnati. June 1953. 27p. diagrs., 6 abs.
(NACA TN 2958)
The vapor-phase oxidation of isooctane at 5000 C
under conditions leading to rapid quenching of the
reaction yields hydrogen peroxide, daisobutylene,
and isobutylene as the major reaction products. As
the reaction time increases, the formation of acetone
and formaldehyde becomes of primary importance.
Under otherwise similar conditions, n-heptane is
attacked at 3500 C to yield a mixture o organic
peroxides as the major initial product. The next
phase of reaction develops with extreme rapidity,
leading mainly to the formation of a mixture of
aidehydes and ketones (principally formaldehyde).
The marked differences between the oxidation be-
havior of these t~wo hydrocarbons are interpreted on
the basis of the temperature required for oxidatrve
attack and of the thermal stabillrty of the alkyl and
peroxy radicals obtained. Prelimmnary results on
>for ad itona evde for ths enh lia oe.
NACA TN 2959
THEORETICAL INVESTIGATION OF THE SUIPER-
SONIC LIFT AND DRAG OF THIN, SWEPTBACK
WINGS WITH INCREASED SWEEVLP .NEAR THE
ROOT. Doris Cohen and Mlorris D. Friedman.
June 1953. 51p. diagrs. (NACA TN 2959)
Formulas are derived by the use of linear theory
for the lift and drag due to lift, at supersonic speeds,
of thin flat wingse having a discontinuity in the
leaging-edge weep, with the inboard portion of the
leading edge very highly swept, the outboard portion
less so, Examples show the effect of the bend in the
leading edge on the pressure distribution, the Hlit-
curve slope, and drag. The results are also related
to the effect of a wing-fuselage juncture on the lift
on the wing;.
A review is presented of available information on the
behavior of britte and ductile materials under condi-
tions of thermal stress and thermal shock. For
brittle materials, simple formula relating physteal
properties to thermal-shock resistance are derived
and used to determine the relative sagnruicance of
two indices currently in use for rating materials.
The importance of simulating operating conditions in
thermal-shock testing is deduced from the formula
and is exrperimentally illustrated by showin that
BeO could be both inferior or superior to A1203 in
thermal shock depending on the testing conditions,
For.ductile materials, thermnal-shock resistance
depends upon the complex interrelation among
several metallurgacal variables which seriously
affect strength and ductility. These variables are
briefly discussed and illustrated from literature
sources. The importance of simulating operating
conditions in tests for rating ductile materials is
especially to be emphasized because of the
importance of testing conditions in metallurgy. A
number of practical methods that have been used to
minimize the deleterious effects of thermal stress
and thermal abock are outlined.
NACA TN 2954
THE STRUCTURE OF TURBULENCE IN FULLY
DEVELOPED PIPE FLOW. John Lanter, National
Bureau of Standards. June 1953. 53ap. diagrs.
(NACA TN 2954)
Measurements, principally with a hot-wire anemo-
meter, were made in fully developed turbulent flow
in a 10-inch pipe at speeds of 10 and 100 feet per
second. It is shown that rates of turbulent-energy
production, dissipation, and diffusion have sharp
maximums near the edge of the laminar sublayer and
that there exist a strong movement of kinetic energy
away from this point and an equally strong movement
of pressure energy toikrard it. It is suggested that the
flow field may be divided into three regions: Wall
proximity where turbulence production, transfer,
and viscous action are of about equal importance; the
central region of the pipe where energy diffusion
predominates; and the intermediate region where the
loca rate of chapg of ur a en -eneg iprdu ion
NACA TN 2056
CRE~EP-BUCKLING ANALYSIS OF RECTANGULAR-
SECTION COLUMNS. Charles Libove. June 1953.
24p. diagrs. (NACA TN 2956)
A previous analysis of the creep behavior of a
slightly served pin-ended H-section column undr
constant load is extended to the slightly curved solid
rectangular-section column. The analysis leads to
a differential equation for the plastic strains at the
midheight cross section. The.form of the equation
indicates the significant parameters which may be
useful in plotting test data on the creep life of col-
umns. These are a lifetime parameter t'cr, an
initial-straightness parameter 8 or 5', and the
ratio of the average applied stress to the Ealer
stress tr/oE. A nuerical method of solving the
RESEARCH ABSTRACTS NO.45
NACA TN 2960
DRAG OF CIRC ULAR CYLINDERS FOR A WIDE
RANGE OF REYNOLDS NUMBERS AND MACH
NUMBERS. Forrest E. Gowren and Edward W.
Perkins. June 1953. 36p.diagrs., photos.
(NACA TN 2960. Formerly RM A52C20)
Pressure distribution measurements on a two-
dimensional circular eylindr have been made at
high subsonic and supersonic speeds. Drag coeffi-
clents obtamned from these measurements are
presented with results from other sources. It was
found that a maximum drag coefficient of about 2. 1
occurs near sonic velocity. As the Mach numer
was increased to 2. 9 the drag coefficient decreased
to about 1. 34. No effects of Reynolds numer were
fomid at supercratical Mach numbers. Effects of
fineness ratio on drag of three-dimensional cylinders
at supersonic speeds were investigated and found to
NACA TN 2961
SUBSONIC FLOW OF AIR THROUGH A SINGLE-
STAGE AND A SEVEN-STAGE COMPRESSOR.
Chung-Hua Wu. June 1953. 32p. diagrs., 2 tabs"
(NACA TN 2961)
A method recently developed for solving the steady
flow of a nonviscous compressible fluid along a
relative stream surface between two adjacent blades
in a turbomachine is applied to investigate the low-
and high-speed subsonic air flow through a single
and through a seven-atage axial compressor.
Appreciable radial flowr results from the radially
increasing values of the angular momentum of the
air particles associated with the Ysymmetrical
velocity diagram at all radiiD used in the exaples.
Tlhe axial-velocity distribution obtained for the .
example checked well with a simple approximation
solution obtained previously by assuming simple
sinusoidal radial-flow paths.
NACA TN 2962
EFFECT OF ICE AND FROST FORMATIONS ON
DRAG OF NACA 651-212 AIRFOIL FOR VARIOUS
MLODES OF" THERMAL ICE PROTECTION. Vernon
H. Gray and Ulwe H. von Glahn. June 1953. 68p.
diagrs., photos. (NACA TN 2962)
Studies wvere mad to determine the effect of ice and
frost formations on the drag of an 8-foot-chord
NACA 651-212 airfoil. At high angles of attack (8oy
glaze-ice formations on the upper surface near the
leading edge of an airfoil caused large increases in
drag and incipient stalling of the airfoil. Runhack
icing on the lower surface, except for heavy span-
wrise ice ridges, presented no serious drag problems.
Rime-ice formations on the leading edge did not
cause large drag increases. Cyclic de-icing of the
leading edge successfully decreased the drag almost
to the bare airfoil drag value. Frost formations on
airfoil surfaces caused large drag increases and
may result in sting of the airfoil.
NACA TN 2963
EFFECT OF VARTTIO~N IN VT STRENGTH ON
THIE AVERAGE STRESS AT MAXIUM LOAD FOR
ALUMINUM-ALLOY, FLA, Z-STIFFENED)
COMPRESSION PANELS THAT FAIL BY LOCAL
BUCKLING. Norris F. Dow, William A. Hickman
and B. Walter Rosen. June 1953. 17~p. diagrs.,
2 tabs. (NACA TN 2963)
A study is made of the effect of variation in rivet
strength on the average stress at maximum load for
7558-T6 aluminum-alloy, flat, E-stiffened compres-
sion panels that fail by local buckling. A curve is
presented for the determination of the relationship
between strength, diameter, and pitch of the rivets
and the strength of stiffened, flat compression
panels of such proportions that failure is by local
NACA TN 2964
A REVISED FORMULA FOR THE CALCULATION
OF GUST LOADS. Kermit G. Pratt. June 1953.
15p. diagrs. (NACA TN 264)
A revised gust-load formula with a new gust factor is
derived to replace the gust-load formula widely used
in gust studies. The revised formula utilizes the
same principles and retains the same simple form of
the original formula. The gust factor is calculated on
the basis of a one-minus-cosine gust shape and is
presented as a function of mass ratio as compared to
the ramp gust shape and wing loading, respectively,
used for the alleviation factor.
NACA TN 2968
PROPELLER-NOISE CHARTS FOR TRANSPORT
AIRPLANES. Harvey H. Hubbard. June 1953. 47p.
diagrs., photos., tab.` (NACA TN 2968)
Calculations of rotational-noise and vortex-noise
levels at a distance of 300 feet for engine ratings of
1, 000 to 10, 000 horsepower have been made for a
large number of propellers in static operation.
Propellers with three, four, six, and eight blades
and diamneters of 8, 12, 16, and 20 feet are
considered. The results are presented in chart and
table form for rapid estimation of the noise levels
and spectrums in the range of tip Mach numbers from
0. 40 to 1. 00. Applications of the data to tandem, and
dual-rotating configurations are given and the
supersonic-type propeller is also briefly considered.
NACA TN 2969
THE CONDENSATION LINE OF AIR AN THE
HEATS OF VAPORIZATION OF OXYGEN AND
NITROGEN. George T. Furukawa and Robert E.
McCoskey, National Bureau of Standards. June
1953. 30p. diagrs., 8 tabs. (NACA TN 2969)
RESEARCH ABSTRACTS NO.45
lation wvas found to exrist between number of grain
per blade cross section and blade life for the ex-
perimentally cast blades; longer life was associated
with the coarser grain sizes. Initial failure time
and uniformity of life was better for the experunen-
tally cast blades poured at the higher nominal sem-
perature than those poured at the lower nominal tef-
perature. The experimentally cast coarse-grained
blades had better uniformity, similar initial failure
time, but lower mean life than a group of commer-
cially cast coarse-grain blades.
NACA RMI E53D23
PROCEDURE FOR MEASURN LIQUID-WATER
CONTENT AND DROPLIET SIZES IN SUPERCOOLED
CLOUDS BY ROTATING MULTICYLINDER METHOII
William Lewis, Porter J. Perkins and Rinaldo J.
Brun. Appendix C: ALTERNATE METHODOF
REUCING ROTATING MULTICYLINDER DATA.
Paul T. Hacker. June 1953. 48p. diagrs., photos.,
4 tabs. (NACA RM E53D23)
The rotating multicylinder method for in-flzght deter-
mination of liquid-wrater content, droplet size, and
droplet-size distribution in icing clouds is described.
Tlhe theory of operation, the apparatus required, the
technique of obtaining data in flight, and detailed
methods of calculating the results, merluding neces-
sary charts and tables, are presented.
NACA RM L53EO4a
PRELIMINARY INVESTIGATIONS OF STRENGTH
CHARACTERISTICS OF STRUCTURE ELEMENTS
AT ELEVATED TEMPERATURES. Eldon E.
Mathauser and Charles Libove. June 1953. 12p.
diagrs. (NACA RM L53EO4a)
Methods for predicting structural strength of col-
umns, plates, and stiffened panels at uniform elevated
temperature are surveyed. For columns and plates,
the material stress-strain curve is needed in deter-
mining short-time strength where creep effects are
negligible, and, for stiffened panels, an "effective"
stress-strain curve is useful. A method for deter-
mining the long-time, or creep, strength of columns
is given. Some data obtained from, creep tests on
stiffened panels are included.
Department of Scientific and Industrial Research
(Gt. Brit. ) RESULTS OF EXLPERIMENTS ON
DISC-TESTERS WITH ROLLERS OF BALL-
B]EARING STEEL AND 'NITRODUR' IN LINEAR
CONTACT WITH HARENED TEST-PIECES.
(Erfahrungen mit rollen aus kugellagerstahl and
Nitrodur bei den versuchen an den wakzenpr'ieftiimden
mit linienberfibrung an gehdrteten priiflingen. ) K.
-W. Kraupner. 10p. diagrs., photos. (DSIR
Sponsored Research (Germany) Rept. 20. Trans.
from Technische Hochschule Braunschweig,
Institute fhzr Maschinenelemente, Einzelbericht 132,
October 27, 1950)
The condensation pressure of air was determined
over the range of temperature from 60o to 850 K.
The experimental results were slightly higher than
the calculated values based on the ideal solution law.
H~eat of vaporization of oxygen was determined at
four temperatures ranging from about 68o to 010 K
and of nitrogen similarly at four temperatures
ranging from. 62o to 78o Kt.
NACA TN 2970
PHYSICAL PROPIERTIrES OF CONCENTRATED
NIRIC ACID, W. L. Sibbitt, C. R. St. Clair, T`. R.
Bump, P. F. Pagerey, J. P. Kern and D. W. Fyfe,
Purdue University. June 1953. 19p. diagrs.,
2 tabs. (NACA TN 2970)
From a review of the literature and additional
experimental measurements, recommended values
of the physical properties of white fuming nitric acid
were determined for the temperature range of
approximately -350 to 3000 F. These properties
included thermal conductivity, dynamic viscosity,
specific heat, density, and vapor pressure. Values
were determined within an accuracy of 5 percent for
all physical properties except the vapor pressure,
NACA RM E53C27
COMPARSO OF SEVERAL METHODS OF CYCLIC
DE-ICING OF A GAS-HEATED AIFOIL. Vernon
H. Gray and Dean T. Bowden. June 1953. 66p.
diagrs., photos., 2 tabs. (NACA RM E53C27)
Several methods of cyclic de-icing of a gas-heated
airfoil were investigated to determine ice-removal
characteristics and heating requirements. The cyclic
de-icing system with a spanwise ice-free parting
strip in the stagnation region and a constant-
temperature gas-supply duct gave the quickest and
most reliable ice removal. Heating requirements
for the several methods of cyclic de-icing are
compared, and the savings over continuous ice
prevention are shown. Data are presented to show
the relation of surface temperature, rate of surface
heating, and heating time to the removal of ice.
NACA RM E53DO6
INVESTIGATION OF EFFECTS OF GRAIN SIZE
UPON ]ENGN LIFE OF CAST AMS 5385 GAS TUR-
BINE BLADES. C. A. Hoffman and C. A. Gyorgak.
July 1953. 21p. diagr., photos., 5 tabs. (NACA
An investigation was conducted to (a) determine the
effects of pouring temperature and grain size upon
uniformity of lives and initial failure times of groups
of experimentally cast AMS 5385 gas turbine blades
aged 24 hours at 13500 F and (b) relate individual
lives of these experimentally cast blades to grain
size. Grain size varied from 4 to 11, 200 (A. S. T. M.
8) grains per cross section. Commercially pro-
duced AMIS 5385 blades were used as a control
group. The blades were operated in a small gas
turbine at a temperature of 14500 F and a midspan
stress of 19, 000 pounds per square inch. A corre.
RESEARCH ABSTRACTS NO.45
This report presents the results of experiments with
normal rollers of ball-bearing steel in linear
contact and experiments wnh pressure rollers of
'Nitrodur 81. '
Royal Aircraft Establishment (Gt. Brit.)
THE ESTIMATION OF ATMOSPHERICS REFRACTION
FROM AZIMIUTH AND ELEVATION DATA. D. J.
Richardson. F broary 1953. 13p. diagram. (RAE
A method is proposed for estimating atmospheric
refraction from azimuth and elevation data
simultaneously with the determination of position
coordinates, taking the enrvature of the earth also
Royal Aircraft Establishment (Gt. Brit.)
CABIN AIR CONDPITIONIN IN HIG SPEED
FLIGHT. F. V. Davies. January 1953. 31p.
diagrs., 3 tabs. (RAE Tech. Note Aero 2196)
Some heat transfer problems of air condiioning at
high speeds are discussed and tentative requirements
given for 0. 5< M< 2. 0. A rational formulation of
requireent cannt be obtained until there is a
background of experimental data on heat transfer and
temperature distributions within the cabin itself and
for the thermal resistance of the cabin walls.
Cooling requirements become more severe as speeds
increase and there are indications that unless the
heat flow through the walls can be considerably
reduced by insulation a cooling system supplementary
to the air conditioning plant may be necessary for
exceptional flight conditions.
Royal Aircraft Establishent (Gt. Brit )
D). V. ]L. HIGH SPEED WIND TUNNEL TESTS ON A
CUSPED TRAILING EDGE AEROFORL NACA 0 00
1 0. 5550/0. 5. F. N. Kirk. January 1953. 18p.
diagram. (RAE Tech.Note Aero 2216)
This note presents a few results from German high
speed win tunnel tests made, in 1945, on a cusped
trailing edge symmetrical section. The results of
the tests on the corresponding airfoil (chord =
0. 850 m) with the full trailing edge angle are not
available and no direct comparison is possible.
Some of the effects of ensping, however, are shown
by comparison wrrith results of tests of a 0. 5 m chord
of the full trailing edge airfoil.
Aeronautical Research Council (Gt. Brit.)
NOTE ON THlE BOUNDARY LAYER NEAR A STAG-
NATION PINT. J. C. Cooke. June 18, 1952. 4p.
2 tabs. (ARC 15,016; FIM 1742)
Homnann's equation for the symmetrical flow of fluid
over a body of revolution near the stagnation point is
shown to be reducible to a well-known equation of
Falkner and Skan, solved by H~artree. The numeri-
cal solution aHom-lolann's equation is given here to
four decimal places as a comparison with the solu-
tion obtained directly bty Miss Hanah
A~ronau cORe Earc oi (t Bi.OB
ROTA~TING DSCS. K.Stewartson. June 30, 1952.
Op. diagrs. (ARC 15,035; FM 1747)
The steady motion of a viscous fluid confined between
two coaxial rotating disks is discussed both experi-
mentally and theoretically. It, ts found experimen-
tally that when the disks rotate in the same sense
the main body of the fluid rotates as well, but if they
rotate in opposite senses the main body of the fluid
is almost at rest. An adequate theory is found to
National Gas Turbine Establishment (Gt. Brit.)
THE BURNING OF SINGLE DROPS OF FUEL.
PART IV. THE FLOW OF HEAT AND CARBON RES-
ID)UE FORMATION IN DROPS OF FUEL. G. A. E.
Godsave. October 1952. 30p.diagrs., 3 tbs.
(NGTE R. 125)
Investigations have been made of the flow of heat in
single drops of evaporating and burning fuel.
Curves are given showing the distribution of tem-
perature with time within these single drops. Data
for the density, heat capacity, and thermal con-
ductivity of petroleum products are presented in
such a manner that the generalized temperature
distribution curves can? be applied to drops of any
petroleum product, given only the normally sgeci-
fied value of the specific gravity at 600 F/60 F.
These invesitigfations are of general interest in
connection with the associated studies of the heat
transfer to evaporating and burning single drops of
fuel. They have also added significance in that cal-
culations based on the results enable conclusions to
be drawn concerning the detailed mechanism of car-
bon residue formation during the combustion of drops
of heavy fuels.
RESEARCH ABSTRACTS NO.45
DIFIFERENTIAL EQUATIONS OF GAS TURBINES
WITH INFINITE NUMBER OF BLADES AND THEIR
INTEGRALS. (Differentsial'nve uravnenlya
gazovykh turbin a beaknonechno bol'shim cyslam
lopastei iikh integraly.) V. M. Astafyev. 6').
(Trans. frolm AkademiiB Nank SBSR, Doklady,
vol. 68, No. 3, 1949, p. 449-452.)
The schematization of the flow in turbines by re-
placing the finite number of blades by an abst ract
system with infinite number of thin blades was
considered as early as 1905-07 by Stodola. In the
present paper, a new variant is given of the
representation of this abstract model which leads
to new results.
DIECLASSIFIEID NACA REPO RTS
NACA RM A6G22
A T~UDIY OF SEVERAL PARAMETERS CONTROL-
LING THE TRAJECTORIES OF A SUPERSONIC
ANTIAIRCRAFT MISSILE POWERED WITHIN SOLID
OR LIQUID-FUEL ROCKETS. Ralph F.
Huntaberger. April 24, 1947. 42p. diagrs. (NACA
RM A6G22) (Declassified from Restricted, 6 '5/53)
The trajectories for a supersonic aniareraft
missile were calenlated by a step-by-step integra-
tion method for a number of different condtin.
The effects of changing drag, initial thrust ratso,
and weight ratio, which are the principal variables
controlling the trajectory for a fixed launching angle,
NACA RM A7A31
DEVELOPMENT OF NACA SUBMERGED INLETS
AND) A COMPARISON WITrH WIN LEADING-EDGE
INLETS FOR A 1/4-SCALE MODEL OF A FIGHTER
AIRPLANE. Emmet A. Mossman and Donald E.
Gault. August 7, 1947. 41lp. diagrs., photos. 11
tabs. (NACA RM A7A31) (D~eclassified from
Characteristics of NACA suberged duc entries and
wing leading-edge inlets designed for a 1/4l-scale
flow model of a fighter-type airplane powFered by a
jet engine in the fuselage are presented. Duet total-
head losses at the simulated entrance to the jet
engine and pressure distributions over the duct
entries are shown. A comparison of the dynai
pressure recovery and critical Mach number of the
two intake systems is made. Included is a discussion
of methods of ameliorating a duct-flow instability
which may appear with a twin-entrance submerged
NACA Rept, 871
Errata No. 2 on a DETERMINATION OF ELASTIC
STRESSES IN GA-TURBIN DISKS.* 8. S. Manson*
California U., Berkeley.
BOUNARY LAWYER EFFECT ON THE SURFACE
PRESSURE OF AN INFINITE CONE IN SUPERSONIC
FLOW. T. C. Lin, S. A. Schaaf and F. S. Sherman.
March 5, 1951. i, 10p. diagrs. (California U.,
The theory of Taylor and Maccoll gives the surface
pressure on an infinite cone in supersonic flow as a
function of the cone vertex angle and the free stream
Mach number and static pressure for a gas of
vanishing viscosity. When a slender conical probe is
used together with an impact pressure probe to
determine the static pressure and Mach number in a
low density gas stream, it is desirable to have some
theoretical estimate of the effect of viscous boundary
layer on the probe readings. A simple approximation
for a conical probe based on linearized supersonic
flow and compressible boundary-layer theory is
Institute of Engineering Research, U. of Calif.
VISCOUS EFFECTS ON IMPACT PROBES IN A
SUBSONIC RAREFIED GAS FLOW. E. D. Kane
and S. A. Schaaf. 1March 9, 1951. 21p. photos.,
diagrs., tab. (Institute of Engineering Re~search,
U. of Calif. HE-150-82)
The investigations described in this report were
made to demonstrate the viscous effect on impact
pressure measurements and to evaluate it by
determining experimentally the relation between
the measured impact pressure, the local reservoir
pressure, the Mach and the Reynolds number In
subsonic compressible flow.
Mount Washington Observatory*
SUMMARY OF A STUD)Y OF NEW PRICIPLES
APPLICABLE TO THE MEASUREMENT OF
PARAMETERS OF THE ICING CLOUD. D.
Davidson. 1953. 44p. diagrs., 2 tabs. (Mount
This report describes work done on the development
of a simple practical means for measuring icing
parameters in natural clouds. The report includes
descriptions of four devices developed and used at
the Mount Washington Observatory together with the
results of observations made with these devices.
RESEARCH ABSTRACTS NO.45
NACA RM A7D11
MEASUREMENTS OF THE DAMPING IN ROLL OF
LARGE-SCALE SWEPT-FRWAR ANDWEPT-
BACK WINGS. Lynn W. Hunton and Joseph K. Der.
July 30, 1947. 39p. diagrs., photos., 2 tabs.
(NACA RM A7Dll) (Declassified from Restricted
Wind-tunnel tests of five large-scale tapered wings
which had angles of sweep of Oo, lr300, and ?1450 have
been conducted to determine the effects of both scale
and sweep on the damping-in-roll paramneter Cl '
Rolling moment and pressure disribuio were
measured for each plain wing while in steady roll for
an angle-of-attack range of -lo to 29o- The eff** *
of both Reynolds number and deflection of partial"
span split flaps were determined from less com-
prehensive tests. Several methods of predicting both
the damping-in-roll and autorotational
Characteristics of the swept wings have been
analyzed, and predicted results have been compared
with the experimental data.
NACA RM A7F06
AN INVESTIGATION AT LOW SPEED OF A LARiGE-
SCALE TRIANGULAR WING OF ASPECT RATIO
TWO. I CHARACTERISTICS OF A WING HAVING
A DOUBLE-WEDGE AIRFOIL SECTION WT ]MAXI-
MUM THICKNESS AT 20-PERCENT CHORD. Adrien
o. tb ( oA RM A'FO ( eca 1 rom
An investigation has been made at the low-speed
characteristics of a 25-foot span triangular wing
hain an aspect ratio of 2. The airfl section of
the wing was a symmetrical double wedge with 5-
percent maximum thichness at 20-percent chord.
Fore and moment data were obtained at several
angles at sidealip for various configurations of 18.5-
pecn area, costn-chord split flaps, 10-percent-
chord nose flaps, and semispan splt-flap-type ailer-
ons. Lift and drag data were obtained from the plain
wing through a limited angle-of-attack range for
Reynolds nube varing between 13 and 34 million,
as based an the mean aerodynamic chord.
NACA RM A7F30
A SUGGESTED METHOD) OF ANALYZING FOR
TRANSONIC FLUTTER OF CONTROL SURFACES
BASED ON AVAIALE EXPERIMENTAL
EVIDENCE. Altbert L. Erickson and Jack D.
Stephenson. December 16, 1947. 59p. diagrs.,
photos., 2 tabs. (NACA RM ATF30) (Declassified
from Confidential, 6/5/53)
Results of a transonic flutter investigation using
a three-dimensional wig with a 651-213, a = 0. 5
airfail section are analyzed. An empirical theory
is developed and checked against the results of two
additional tests of smal-scale models.
NACA RMI A7G18
WING-FLO TESTS O1F A TRIANULAR WIG OF
ASPECTS RAIO TWO. I. EFFECTIVENESS OF
SEVERAL TYPES OF TRAILINGEGE FLAPS ON
FLA-PLATE MOELS. George A. Rathert, Jr.
and George E. Cooper. November 14, 1947. 61p.
diagra.,photoa. (NACA RM A7Gl8) (Declassified
from Confadential. 6/5/53)
Ths report contains the results of an investigation
conducted by the wing-flow method of the effective-
ness of several different types of longitudinal control
flap on a flat-plate triaglr plan form of aspect
ratio 2. The control effectiveness parameters
da/d~ii and dCm 'ddl were determie for CN=0
from a Mach number of 0. 50 to 1. 10 at Reynolds
numbers of 500, 000 to 1, 300, 000. Neither the basic
plan form nor any of the controls exiie any
critical stick-fixed characteristics in the transonic
NACA RM ATH19
CHARACTERISTICS OF A 15-PERCENT-CHORD
AND A 35-PERCENT-CHORD PLAIN FLAP ON
THE NACA 0006 AIRFORL SECTION AT HIG SUTB-
SONIC SPEEDS. Richard J. Ilk. October 2, 1947.
33p. diagrs., tab. (NACA RMl A7H19) (Declassified
from Restricted, 5/6/53)
Wind-tunnel tests have been made to determie the
aerodynamic characteristics of a 15i-percent- and a
35-percent-chord plamn trailing-edge flap on the
NACA 0006 airfoil section. Simultaneous measure-
ments of section lift, drag, and pitching moment
were made over a range of Mach numbers from 0. 3
to approximately 0. 9 at angles of attack ranging from
_go to 12o for flap deflections of Go, 50, and 100,
Increments of section lift coefficient and changes in
airfoil angle of attack necessary to maintain constant
lift with unit changes in flap deflection are presented
as a measure of the elTectiveness at high subsonic
speeds of a plain flap employed on a very thin profile.
NAA RMI AT28
AN INVESTIGATION AT LOWN SPEED OF A LARGE-
SCALE: TRIANGULAR WING OF ASPECT RAIO
TWO. II THE EFFECT OF AIRFORL SECTION
MODIFICATIONS AND THE DETERMINATION OF
THE WAKE DOWNWAS. Adrien E. Anderson.
December 10, 1947. 78p. diagrs., photos., tab.
(NACA RM ATH28S) (Declassified from Restricted,
This report contained force and moment data obtained
throughout the angle-of-attack range for a triaguar
wing havin a symmetrical double-wedge airfoil see-
tion and various degrees of rounding of the win
leading edge and maximum thickness. The dynamic
pressure and downwash angle in the wake are also
RESEARCH ABSTRACTS NO.45
NACA RM ETB11a
COOLING OFGAS TURBNE. I EFFECTS OF
ADDITION OF FINS TO BLADE TIPS AND ROTOR,
ADMISSION OF COOLING AIR THROUGH PART OF
NOZZLES, AN CHANGE IN THERMAL COkDUC-
TIVITY OF TURBINE~ COMPONENTS. W. Byron4 -
Broawn. February 11, 1947. 26p. diagrs.
(NACA RM E7B11a) (DeclasslZled from Restricted,
An analysis was developed for calculating the radial
temperature distribution in a gas turbine with only
the temperatures of the gas and the cooling air and
the surface heat-transfer coeffacients known. Thla
analysis was applied to determine the temperatures
of a complete wheel of a conventional single-stage
impulse exhaust-gas turbine. The temperatures were
first calculated for the case of the turbine operating
at design conditions of speed, gas flow, etc., and
with only the customary cooling arising from
exposure of the outer blade flange and one face of the
rotor to the air. Calculations were next made for the
case of fins applied to the outer blade flange and the
rotor. Finally the effects of using part of the nozzles
for supplying cooling air and the effects of varying
the metal thermal conductivity from 12 to 260 Btu
per hour per foot per oF on the wheel temperatures
were determined. The gas temperatures at the
nozzle box used in the calculations ranged from
16600 F to 2000o F.
NACA RM ETB11b
COOLING OF GAS TURBINES. II EFFECTIVE-
NESS OF RIM COOLING OF BLADES. Lincoln
Wolfenstein, Gene L. M~eyer and John 8. McCarthy.
March 18, 1947. 31p. diagrs. (NACA RM E7Bllb)
(Declassified from Restricted, 6/5/53)
An analysis is presented of rim cooling of gasi-
turbine blades; that is, reducing the temperature at
the base of the blade (wheel rim), which cools the
blade by conduction alone. Formulas for tempera-
ture and stress distributions along the blade are
derived and, by the use of experimental stress-
rupture data for a typical blade alloy, a relation is
established between blade life timee for rupture),
operating speed, and amount of rim cooling for
several gas temperatures. The effect of a blade
parameter combining the effects of blade dimensions,
blade thermal conductivity, and heat-transfer
coefficient is determined. The effect of radiation on
the results is approximated.
NACA RM E7G23
INVESTIGATION OF THRUST AUGMENTATION OF
A 1600-POUN THRUST CENTRIFUGAL-FLOW-
TYPE TURBOJET ENGINE BY INJECTION OF
REFRIGERANTS AT COMPRESSOR INLETS.
William L. Jones and Harry W. Dowman. August 25,
1947. 44p.diagrs., photo., tab. (NACA RM E7G23)
(Declassified from Restricted, 6 '5 53)
NACA RM A7I29
TESTS OF SUBMlERGED DUCT INSTALLATION ON
A MODIFIED ]FIGHTER AIRPLANE IN THE. AMES
40- BY 80-FOOT WIND TUNNL. Norman J-
Martin. December 11, 1947. 25p. diagrs., photos.,
2 tabs. (NACA RM A7129) (Declassified from
An investigation of an NAA submerged intake! instal-
lation on a modified fighter airplane was conducted to
determine the full-scale aerodynamic characteris-
ties of this installation. In addition, tests were
conducted on the submerged inlet with revised
entrance lips and deflectors to determine the con-
figuration which would result in the best dynamic
pressure recovery measured at the inlet for this
installation without a major rework of the entrance.
NACA RM A7IL11
LOW-SPEED INVESTIGATION OF A SMALL TRIAN-
GULAR WING OF ASPECT RAI 2.0. II FLAPS.
Leonard M. Rose. Agut 9, 1948. 17lp. diagrs.,
photo. (NIACA RM ATL11) (Declassified from
Split and plain trailing-edge flaps were tested on flat
plate models having an aspect ratio of 2.0 and 63.4o
sweepback of the leading edge. A constant chord
flap was found to be the most effective and a skewed
wing tip flap the least effective. Some additional
tests of short-span split flaps at several chordwise
locations indicate the possibility of obtaining
balancing moments without loss in lift or changes in
lift without change in attitude or balance.
NACA RM A8CO3
LOWT-8PEED INVESTIGATION OF A SMALL TRIAN-
GULAR WING OFr ASPECT RATIO 2.0. III STATIC
STABILITY WITH TWIN VERTICAL FINS. Leonard
M. Rose. August 24, 1948. 11p. diagrs. (NACA
RM A8CO3) (Declassified from Restricted, 6/11/53)
Twin vertical fins were tested on a triangular wing.
It was found that these fins maintained substantially
constant effectiveness throughout the angle-of-attack
range. The addition of the twin fins resulted in a
loss in maximum lift and a rednetion in static longi-
tudinal stability at lift coefficients above 0.4.
NACA RM A9126
ESTIMATION OF THE FORCES AND MOMENTS
ACTING ON INCLINED BODIES OF REVOLUTION OF
HIGH FINENESS RAIO H. Juin Allen.
November 14, 1949. 27p. diagrs. (NACA RM A9126)
(Declassified from Restricted, 8/11/53)
An approximate theory including an allowance of the
effects of viscosity on inclined bodies of revolution is
developed. Calculations based on this theory are
compared with experiment. Discussion of the proba-
ble effects of Reynolds and Mach number is included.
RESEARCH ABSTRACTS NO.45
The performance of a centrifugal-flow -type turbojet
engine (having a normal military rating of 1600-16
thrust at a rotor speed of 16,500 rpm), has been in-
vestigated at zern flight speed with injection of
refrigerants at the compressor inlets. The largest
part of these investigations was devoted to the injee-
tion of water and waer-alohol mixtures; brief in-
vestigations were also conducted with the injection of
kerosene and carbon dioxide.
NACA RM E7G29
FLIGHT INVESTIGATIONN OF THRUST AUGMIENTA-
TION OF A TURBOJET ENGINE BY WYATER-
ALCOHOL INJECTION. Carl Ellisman.
September 29, 1947. 20p. diagrs., photos. (NACA
RM E7G29) (Declassified from Restricted, 6/5/53)
Thrust augmentation by injection of water-alcohol
mixtures into compressor inlets of turbojet engines
was investigated in flight at sea level, 5, 000 and
10,000 feet to determine water-alcohol mixture and
injection rate for optimum thrust augmentation. The
mixture of 20 percent alcohol by weight and an injec-
tion rate of 1.45 pounds per second produced maxi-
mum thrust augmentation at 10,000 feet. This injec-
tion rate was equivalent to a water-alcohol-to-air
ratio of approximately 0.05 and gave a 21 percent in-
crease in net thrust. Below 10,000 Ifeet the optimum
water-alcohot-to-air~ ratio was within the range of
0. 03 to 0l. 05.
NACA RM E7G30
COOLING OF GAS TURBINES. VII EFFECTIVE-
NESS OF AIR COOLING OF HOLLOW TURBINE
BLADES WTITH INSERTS. Joseph R. Bressman and
John N. B Livingood. October 20, 1947. 51p.
diagrs., 2 tabs. (NACA RM E7G30. Now Rept. 094)
(Declasstfied from Restricted, 6/5/53)
Contamns results of analytical investigation to deter-
mine reduction in cooling-air requirements and
increase in effective gas temperature for same quan-
tity of cooling air resulting from use of insert in
cooln .ai pas age ofehollow a r- 1 dddturbineed
conduction on dilution and effects of variations in
blade-root, average insert, and blade-root cooling-
air temperatures on blade temperature are included.
Insert that reduced cooling-air mass flow to minimum
hot-gas flow ratio one-half reduced air mass flow
two-thirds and three-fifths for adequate cooling.
NACA RMI E7I22a
COOLING OF GAS TURBINES. VIII THEORETICAL
TEMPERATURE DISTRIBUTIONS THROUGH GAS
TURBINE WITH SPECIAL BLADES AND COOLING
FINS ON THE RIM(. W. Byron Brown and John
N. B. Livingood. February 17, 1948. 21p, diagrs.
(NACA RM E7I22a) (D~eclassified from Restricted,
Presents effects on blade life and effective gas temr-
perature of analytical investigation of addition of
cooling fins on the rim. Also contains two-
dimensional temperature distributions through the
turbine and three-dimensional temperature distribu-
tions through a section of the rim near a blade root,
both obtained by relaxation. Addition of cooling fins
on the rim doubled blade life and permitted a 300 F
increase in effective gas temperature. Proof was
obtained that a one-dimensional distribution is
suffciently accurate for most applications.
NACA RM E7I25
A COMPARISON OF THIE SIMULATED-ALTITUD)E
PERFORMANCE OF TWO TURBOJET COMBUSTOR
TYPES. Ray E. Bols, Thomas T. Schroeter and
Engene V. Zettle. October 7, 1948. 34lp. diagrs.
(N~ACA RM E7I25) (Declassified from Restricted,
Performance of German Jumo 004 can-type combus-
tor and of two contemporary turbojet combustors of
U. S. design, anuar and can type, were compared to
determine whether advance stage in combustor design
had been reached with Jumo 004 and to determine
whether there are basic, inherent differences in per-
formance achieved with either can or annular type.
Combustors are compared with respect to altitude
operational limits, combustion efficiency, and total-
pressure! loss across combustors. Neither of two
U. S. combustor types showed basic inherent advan-
tages or disadvantages when compared. The German
combustor under given conditions of investigation
resulted in performance that is generally poorer than
that of U. S. combustors.
NACA RM L6H28a
INVESTIGATION OF THE CHARACTERISTICS OF A
HIGH-ASPECT-RATIO WING IN THE LANGLEY 8-
FOOT HIGH-SPEED TUNNEL, Richard T. Whitcomb.
August 28, 1946. 77p. diagrs., photos., 3 tabs.
(NACA RM L6H26a) (Declassified from Restricted,
An in estig on of t e cha atritc 1 a wigwthh
section has been made at Mach numers up to 0.925.
The wing tested has a taper ratio of 2.5:1.0, no twist,
dihedral, or sweepback, and 20-percent-chord 37.5-
percent-semispan plain ailerons. Results presented
include the normal-force, span-loading, pitching-
moment, drag, and wake-width data for the wing alone
with undeflected straight-sided ailerons.
NACA RM L6H28b
PRELIMINARY INVESATIGTON OF DOWNASH
FLUCTUATIONS OF A HIGH-ASPECT-RATIO WING
IN THE LANGLEY 8-FOOT HIGH-SPEED TUNNEL.
Antonio Ferri. August 28, 1946. 25p. diagrs.
(NACA RMn L6H28b) (Declassified from Restricted,
RESEARCH ABSSTRACTS NO.45
A supersonic inlet wit supersonic deceleration of
the flow entirely outside of the inlet is considered.
A particular arrangement with fixed geometry
having a central body wit a circular anuar intk
is analyzed, and it is show theoretacally that this
arrangement gives high pressure recovery igr a
large range of Mach number and mass flowE and 4
therefore is practical for use on supersomec airplanes-
and missiles. For some Mach numbers the drag
coefficient for this type of inlet is larger than the
drag coefficient for the type of inlet with supersonic
compression entirely inside, but the pressure
recovery is larger for all flight conditions. The
differences in drag can be eliminated for the design
NACA RM L6L10a
AN INVESTIGATION OF THIE HINGIE-MOMENfT
FLUCTUATIONS OF 0. 20-CHORD PLAIN AILERONS
ON A HIH-ASPECT-RATIO WING IN TH
LANGLEY 8-FOO HIGHI-SPEED TUNNEL. Arvo
A. Luoma and Luke L. Liccini. January 10, 1947.
9p. diagrs. (NACA RM L6L10a) (Declassified from
Concurrently with a three-dimensional lateral-
control investigation made in the Langley 8-foot high-
speed tunnel of a wing of high aspect ratio having
0.20O-chord, straight-sided-profile plain ailerons, a
few photographic records of aileron vibrations were
obtained by means of an oscillograph connected to the
electrical strain gage used to measure arleron hinge
moments. At supercritical Mnach numbers, frequen-
cies of the order of magnitude of 50 to 100 cycles per
second were observed for the hinge-momnent
fluctuations of the aileron of the model. For a
104. 5-foot-span airplane, full-scale hinge-moment -
fluctuation frequencies (based on the model
frequencies) are indicated to be of the,same order of
magnitude as the wing natural frequencies for an
airplane of this size.
NACA RM L6L10b
INVESTIGATION AT HIGH SPEEDS OF A
HORION A-AI ]MDM EL N THERE LANGBIEat
January 31, 1947. 102p. diagrs., 3 tabs. (NACA
RM L6L10b)(Declassified from Restricted, 6;.5 '53)
Pressure-distribution measurements and elevator
hinge-moment measurements were made to deter-
mine the aerodynaic characteristics of a
horizontal-tail model having an NACA 65-108 airfoil
section equipped with a 30-percent-chord sealed
unbalanced elevator and a 10-percent-chord plain
trim tab. The tests were made for various angles of
attack and control-surface deflections at M~ach numr-
bers ranging from 0. 40 to 0. 90. Data are presented
for tests made with the surfaces of the model smooth
and also with boundry-layer transition fixed at the
0. 10-chord station by a row of carborundum grains
on each surface.
A series of tests has been made in the wak of a
high-speed wing of high aspect ratio at Mach
numbers up to approximately 0. 00 to determine the
fluctuations of the downwash in the zone of possible
tail locations. Serious fluctuations occurred in the
wake of the wring. The fluctuations extended beyond
the wake boundaries but with decreasing amplitude.
For a high angle of attack of the wing and high Mach
numbers the fluctuations wtere very large and extend-
ed as high as approxiately 0. 00 chord above the
chord line of the wing. For medium angles of attack
important fluctuations occurred up to about 0. 70
chord above the chord line of the wing.
NACA RM L6H28c
INVESTIGATIION OF DIVE BRAKES AND A DIVE-
RECOVERY FLAP ON A HIG-ASPECT-RATIO
WING IN TH[E LANGLEY 8-FOOT HIGH-SPEED
TUNNEL. Axel T. Mattson. August 28, 1946.
118p. diagrs., photos., tab. (NACA RM L6H28e)
(Declassified from Restricted, 6/5/53)
T~he results of tests made to determine the aero-
dynamic characteristics of a solid brake, a slotted
brake, and a dive-recovery flap mounted on a high-
aspect-ratio wing at high Mlach numbers are pre-
sented. The data were obtained in the Langley 8-foot
high-speed tunnel for corrected Mach numbers up to
0.040. The results have been analyzed with regard to
the suitability of dive-control devices for a proposed
high-speed airplane in limiting the airplane terminal
Mach number by the use of dive brakes and in
achieving favorable dive-recovery characteristics by
the use of a dive-recovery flap.
NACA RM L6H28d
AN INVESTIGATION OF A HIGH-ASPECT-RATIO
WING RAVIG 0.20-CHORD PLAIN AILERONS IN
THE LANGLEY 8-FOOTr HIGH-SPEED TUNNEL.
Arvo A. Luoma. August 28, 1946. 124p. diagrs.,
photos., 2 tabs. (NACA RIM L6H28d) (Declassified
from Restricted, 6/5/53)
A three-diensional lateral-control investigation was
hain 0,chordsp ofhti dhe poir e pas n e er-
Spanwise loadings and moments and rolling-momlent
coefficients were obtained from pressure-distribution
measurements, and hinge-moment data were ob-
tained by an electrical strai gage for Mach numbers
up to 0.025 at aileron deflections from -10o to 100
and at various angles of attack.
NACA RM L6J31
PRELIMINARY INVESTIGATION OF A NEW TYPE
OF SUPERSONIC INLET. Antonio Ferri and Louis
Mi. Nucci. November 27, 1946. 35p. diagrs.,
photos. (NACA RM L6J31. Now TN 2286)
(Declassified from Restricted, 6/5/53)
RESEARCH ABSTRACTS NO.45
NACA RM L6Ll6
EFFECTS OF COMPRESSIBLITY ON THE CHA-
ACTERISTICS OF FIVE AIRFOILS. Bernard N.
Daley. April 25, 1947. 73p. dlagrs., photos., tab.
(NACA RM L6L16) (Declassified from Restricted,
6 5, 53)
Pressu re -dist ribut ion tests were made for deter-
mination of the effects of compressibility on the
characteristics of the following 5-inch-chord
airfolls: the NACA 66, 2-215, the NACA 66, 2-0:5'
and the NACA 65(216)-418 sectsons representing the
low-drag sections- the NACA 16-212 section, typical
of the type designed for high critical speeds; and the
NACA 23015 section, one of the older conventional
airfoils. Schlieren photographs of the air flow and
limited data concerning the wake characteristics
were also obtained for the conventional airfoil. Data
are presented for an approximate Mach number
range from 0. 34 to 0. 75; the corresponding
Reynolds number range Is from 700, 000 to
1, 800, 000.
NACA RM L7B27
INITIAL TEST IN THE TRANSONIC RANGE OF FOUR
FLUTTER AIRFOILS ATTACHED TO A FREELY
FALLING BODY. J. G. Barmby and S. A.
Clevenson. May 5, 1947. 16p. diagrs., photo., 2
tabs. (NACA RM L7B27) (Declasgidfied from
Restricted, 6. 5 53)
Results of the first test in the transonic range of four
flutter airfoils attached to a freely falling body are
reported. Failures of the airfoils were telemetered
and recorded. These airfoils were designed in an
attempt to spread the range of flutter speeds.
Telemeter records show that three of the airfoils
failed at Mach numbers between 0.87 and 0.90 at an
altitude of 14,000 feet, and the fourth airfoil apparent-
ly faded at a Mach number of 1.025 at 1900 feet above
NACA RM L7JO8
INITIAL FLIGHT TEST OF THE NACA FR-1-A, A
VEHICLE FOR TRANSONIC FLUTTER RESEARCH.
Ellwyn E. Angle. 25p. dagrs., photos. (NACA
RM L7JOB) (Declassified from Restricted, 6/11/53)
The first of a series of flutter rockets, designated
the NACA FR-1-A, was successfully launched and
flown at the Pilotless Aircraft Research Station,
Wallops Island, Va. Two identical swrept wings were
tested; the left wing failed at 967 feet per second
(659 miles per hour, Mach number Mi = 0. 89), and the
right wing remained on the model throughout the
entire flight maximumm Mach number of 1.0O). The
exp~erimental value of failure speed obtained at a
Mach number of 0. 89 was 76. 2 percent greater than
the flutter speed obtained using the two-dimensional,
incompressible, unswept-wing flutter theory.
14ACA RM L7K17
FLUTTER INVESTIGATION IN TH TRASOIC
RANGE OF BI AIRFOHS ATTACHED TO THREE
FREELY FALLING BODIES, S. A. Clevenson and
William T. Lauten, Jr. May 6 1948. 32p. diagrs.,
Photos., 2 tabs. (NACA RM L7K17l) (Declassified
from. Restricted, 6/11/53)
Resuls of six flutter airfoils attached to three freely
falling bodies are reported. Four similar airfoils,
twro unswept and two wit 45o sweepback, fluttered at
tanssonic Mach nmes. Flutter fre tnces andh
the top Mlach nuber of the bomb, M = 1.145, wit-
out flttr or failure. Th results compare favorably
with previous bomb and rocket tests.
NAC RM L8C24
PLIGH~T TEST OF NACA FR-1-B, A LOW-
ACCELERATIO ROCKET-PROPELLED VEHICLE
FOR TRANSONIC FLUTTER RESEARCH., E11wyn E.
Angl, Shermnan A. Clevenson and Reginald R.
Lundstrom. July 20, 1948. 22p. diagrs., photos.,
3 tabs. (NACGA RM L8C24) (Declassified fro
A technique has been developed using a low accelera-
tion rocket with strain-gage-type telemeter for
determining the quantitative flutter chaacteristics of
PM i re di team. ew o50swp Ain ke
fluttered in a symmetrical mode at a frequency of 37
cycles per second at a Mach number of 0.65.
NACA RM L9LO9
THE EFFECT OF CHANGES IN THE LEADING-
EDGE RADIUS ON THE AERODYNAMIC CHARAC-
TERTSICS OF A SYMMETRICAL, 9-PERCENT-
THICK AIRFOIL AT HIGH-SUBSONIC MACH NUM-
BERS. Miltton D. Humphreys and Raymond A.
Robinson. August 7, 1950. 48p. diagrs., photos.,
tab. (NACA RM L9LO9) (Declassified from
Wind-tunnel tests of the NACA 0009-64, 0000-54,
and 0000-44 airfoils indicated that their leading-edge
radii did not greatly influence their high-speed aero-
dynamic characteristics for Mach numbers up to
0. 825. At a Mach number of 0. 85 the drag of the
NACA 0009-54 airfoil, having a leading-edge radius
of 0i.620 percent chord, appeared to be 12 to 25
percent lower than that of the other airfoils for
normal-force coefficients in the range 0 to 0i. 3.
NACA RM L53B18
A METHOD FOR CALCULATING THE AERODY-
NAMIIC LOADING ONWING-TIP-TANK COMBI-
NATIONS IN SUBSONIC FLOW. Samuel W.
Robinson, Jr. and Martin Elotnick. April 7, 1953.
43p. diagrs. (NACA RM L53B18) (Declassified
from Restricted, 4/9/53)
An analytical method for calculating the aerodynamic
loading on wing-tip-tank combinations in subsonic
flow is developed using a simple horseshoe vortex-
image system for the case of the tank axris in the
plane of the wing. An illustrative exape is given
in the appendix, in which wing and tip-tank loadings
are calculated for three configurations. The calcu-
lated results are shown to be in good agreement with
NACA-Lanly- T-10-53 -4000
UNIVERSITY OF FLORIDA
3 1262 08153 300 1
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