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National Advisory Committee for Aeronautics
NOVEMBER 15, 1955
CURRENT NACA REPORTS
NACA Rept. 1175
EFFECT OF VARIABLE VISCOSITY AND THERMAL
CONDUCTIVITY ON HIGH-SPEED SLIP FLOW BE-
TWEEN CONCENTRIC CYLINDERS. T. C. Lin and
R. E. Street, University ol Washington. 1954. Ii,
36p. diagrs. INACA Rept. 1175. Formerly
The differential equations ot slip flo,. including the
Burnett terms, were lrst solid by Schamrberg as-
suning that the coellicients of 'iscositv and heat con-
duction ol the gas were constants. The problem is
solved herein for variable coefficients of viscosity
and theriial conductivity u% applying a transforma-
tion leading to an iteration n;ethod. The method,
starting with the solution for constant coelticients,
enables one to approximate tne solution for variable
coefficients ver closely after one or t.o steps.
Satisfactory results are sho.'n to follow froni
Schanberg's solution by using his values of the
constant coefficients multiplied by a constant
factor Ti, leading to what are denoted as the effec-
tive coefficients of viscosity and thernal conduc-
NACA Rept. 1189
THEORETICAL AND EXPERIMENTAL ANALYSIS
OF LOW-DRAG SUPERSONIC INLETS HAVING A
CIRCULAR CROSS SECTION AND A CENTRAL
S BODY AT MACH NUMBERS OF 3.30, 2.75, AND
2.45. Antonio Ferri and Louis M. Nucci. 1954.
ii, 37p. diagrs., photos. (NACA Rept. 1189.
Supersedes RM L8H13)
Contains theoretical and experimental analysis of
circular inlets having a central body at Mach num-
bers of 3.30, 2.75, and 2.45. The inlets have been
designed in order to have low drag and high pressure
recovery. The pressure recoveries obtained are of
the same order of magnitude as those previously
obtained by inlets having very large external drag.
NACA Rept. 1198
A THEORETICAL STUDY OF THE EFFECT OF
FORWARD SPEED ON THE FREE-SPACE SOUND-
PRESSURE FIELD AROUND PROPELLERS. I. E.
Garrick and Charles E. Watkins. 1954. 11, 16p.
diagrs., tab. INACA Rept 1198. Supersedes
The sound-pressure field due to thrust and torque of
a propeller in flight at uniform subsonic speed is
analyzed by use of a distribution of acoustic doublets
located at the propeller disk. The basic element
used to syntnesize the field is the pressure field of a
concentrated force moving undormly at subsonic
speeds, for which an expression generalizing one of
Lamb's for the fixed concentrated force is given
The results can be regarded as an extension of the
work for the static propeller given by Gutin (NACA
TM 1195) for the far pressure field and given by
Hubbard and Regier (NACA Rep 996) for the f.
near the tips of tT"rot~~m g propeller. The extended
formulas are us4d to cp.4glate the sound field for a
specific two-blae propeller operating at constantI
torque for various forward- speed, MaqSh ~nbe s.
NACA Rept. 1209 U PO
DEVELOPMENT OF TURBULENCE-aSURING \
EQUIPMENT. Leslie S. G. Ko,-fiszil Jghn
Hopkh r,_ Liri'ers t,. 15-1.. ,,, Op. ala .'., '"
photos., tLb. INACA Rept. 1209 Supersedre,
TN 2839) '
Hot-' ire turbulence-mrreasuring equiprAipnL-tas-beerr'
detelopea to nmeet the more-sltrngent requirements
involved in the neasurenient ol fluctuafions"Ti7 flow
paranieters 3a superson-r: ,elocili es. The higher
nmean speed neces.itates the resolution ol higher fre-
quen'y conpounenil th an at o v speed, and the rtla-
li\el low' turbulence level present at supersonic
speed nia.es nec.essar,. a. improved noise level for
the equipnmeit. The equipment covers the frequency
rane fron, 2 to 70,000 cycles per second. The
equipiert is a.id pt.tbl to all-purpose turbulence
iorl. Ith inlpro'v! utilllt, and iccurac,' over lhit of
older t'. pes '.i. equipnEi t. Samniple mreasureienits
.ire given to demonnst rate the perlorniance.
NACA RM E55116
SPARK IGNITION OF FLOWING GASES V AP-
PLICATION OF FUEL-AIR-RATIO AND INITIAL-
TEMPERATURE DATA TO IGNITION THEORY.
Clyde C. Svett. Jr. November 1955 19p. diagrs
INACA RM E551161
Data showing the effect of fuel-air ratio and initial
temperature on spark-ignition energy are presented
and applied to a previously developed theory of igni-
tion. The initial-temperature data are consistent
with the theory; fuel-air-ratio data are only par-
Stally consistent Probable reasons for the discrep-
ancy are discussed
"AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA 1512 H ST NW, WASHINGTON 25, D C. CITING CODE NUMBER ABOVE EACH TITLE;
.THE REPORT TITLE AND AUTHOR
NACA TM 1388
GENERAL SOLUTIONS OF OPTIMUM PROBLEMS
IN NONSTATIONARY FLIGHT. (Soluzioni General
di Problem di Ottimo in Volo Non-Stazionario).
Angelo Miele. October 1955. 25p. diagrs., tab.
(NACA TM 1388. Trans. from L'Aerotecnica, v.32,
no.3, 1952, p.135-142)
A general method concerning optimum problems in
nonstationary flight is developed and discussed.
Various conditions of flight in a vertical plane
(climb with minimum time, climb with minimum fuel
consumption, steepest climb, descending and gliding
flight with maximum time of space) are studied;
the corresponding best techniques of flight, that is,
the optimum speed-height relationships, are deter-
NACA TN 3415
A UNIVERSAL COLUMN FORMULA FOR LOAD AT
WHICH YIELDING STARTS. L. H. Donnell and V. C.
Tsien, Illinois Institute of Technology. October
1955. 48p. diagrs., photos., tab. (NACA TN 3415)
An analysis is presented of the load at which yielding
first occurs in actual columns, taking adequately into
account all the factors which have an important ef-
fect upon this load. The results are expressed as a
formula or chart applicable to all cases.
NACA TN 3476
CALCULATED SPANWISE LIFT DISTRIBUTIONS
AND AERODYNAMIC INFLUENCE COEFFICIENTS
FOR SWEPT WINGS IN SUBSONIC FLOW.
Franklin W. Diederich and Martin Zlotnick.
October 1955. 173p. diagrs., tabs. (NACA
Spanwise lift distributions have been calculated for
61 swept wings with various aspect ratios and taper
ratios and with a variety of angle-of-attack or twist
distributions, including flap and aileron deflections,
by means of the Weissinger method with eight control
points on the semispan. Also calculated were aero-
dynamic influence coefficients which pertain to a cer-
tain definite set of stations along the span.
NACA TN 3494
SOUND PROPAGATION INTO THE SHADOW ZONE
IN A TEMPERATURE-STRATIFIED ATMOSPHERE
ABOVE A PLANE BOUNDARY. David C. Pridmore-
Brown and Uno Ingard, Massachusetts Institute of
Technology. October 1955. 57p. diagrs., photo.
(NACA TN 3494)
A theoretical and experimental study of the sound
field about a point source over a plane boundary in
the presence of a vertical temperature gradient has
been made. Methods are presented for analyzing the
effects of temperature gradients on the attenuation of
sound in the shadow zone of a sound field.
RESEARCH ABSTRACTS NO.92
NACA TN 3495
FAILURE OF MATERIALS UNDER COMBINED RE-
PEATED STRESSES WITH SUPERIMPOSED STATIC
STRESSES. George Sines, University of California
at Los Angeles. November 1955. 69p. diagrs.,
photos., tabs. (NACA TN 3495)
Experiments on blaxial alternating stresses and
simple combinations of static stress with alternating
stress are reviewed. A general criterion for the
effect of static stress on the permissible amplitude
of alternating stress is proposed and compared with
results of tests performed under more complex
stress states. Tests were performed to determine
the effect of static compression on alternating tor-
sion and the results are compared with the general
criterion. A modification of Orowan's theory of
fatigue to include the effect of static stress is pre-
NACA TN 3531
PILOT'S LOSS OF ORIENTATION IN INVERTED
SPINS. Stanley H. Scher. October 1955. 10p.
diagrs., photos. (NACA TN 3531)
The rising problem of pilot orientation during spins,
especially during unintentional inverted spins, is
discussed. The free-spinning-tunnel results and
reported airplane experiences concerning inverted
spins and recoveries are reviewed. Information is
provided regarding the nature of inverted spins,
optimum control technique for recovery, and some
of the factors which apparently contribute to the
pilot's loss of orientation. A spin-simulator rig
which was recently constructed at the Langley Labo-
ratory for use in an attempt to understand better the
problems confronting the pilot of a spinning airplane
NACA TN 3537
HELICOPTER INSTRUMENT FLIGHT AND PRECI-
SION MANEUVERS AS AFFECTED BY CHANGES IN
DAMPING IN ROLL, PITCH, AND YAW. James B.
Whitten, Johr P. Reeder, and Almer D. Crim.
November 1955. 14p. diagrs., photos.
(NACA TN 3537)
The damping in roll, pitch, and yaw of a single-rotor
helicopter was varied by means of electronic compo-
nents, and these variations were evaluated by per-
forming instrument approaches and other precision
maneuvers. Increased damping in roll was found to
be particularly beneficial, whereas corresponding
changes in yaw and pitch were less effective. Some
operational aspects of helicopter instrument ap-
proaches are also included in the discussion.
RESEARCH ABSTRACTS NO. 92
NACA TN 3539
SOME EFFECTS OF SYSTEM NONLINEARITIES IN
THE PROBLEM OF AIRCRAFT FLUTTER.
Donald S. Woolston, Harry L. Runyan, and Thomas
A. Byrdsong. October 1955. 20p. diagrs., tabs.
(NACA TN 3539)
This paper presents the results of a preliminary in-
vestigation of the effect of nonlinear structural
terms on the flutter of a two-degree-of-freedom
system. The three types of nonlinearities investi-
gated were a flat spot, hysteresis, and a cubic
spring. Calculations were made on an analog com-
puter. For one case, the flat spot, an experimental
investigation was also made and good correlation
with theory was found.
NACA TN 3542
ANALYSIS OF STRESSES IN THE PLASTIC RANGE
AROUND A CIRCULAR HOLE IN A PLATE SUB-
JECTED TO UNIAXIAL TENSION. Bernard
Budiansky and Robert J. Vidensek. October 1955.
39p. diagrs., tabs. (NACA TN 35421
An approximate theoretical solution is presented for
the stresses in the plastic range around a circular
hole in an infinite sheet subjected to uniaxial tension.
The solution Is based on the simple deformation
theory of plasticity and is found by application of a
variational principle in conjunction with the Rayleigh-
Ritz procedure and the use of a high-speed computing
machine (SEAC). Numerical results are obtained
for four different materials, which are characterized
by four distinct uniaxial stress-strain curves. The
results for stress concentration factor in the plastic
range are compared with those obtained from a for-
mula due to Stowell.
NACA TN 3544
COMPARISON BETWEEN THEORETICAL AND EX-
PERIMENTAL STRESSES IN CIRCULAR SEMI-
MONOCOQUE CYLINDERS WITH RECTANGULAR
CUTOUTS. Harvey G. McComb, Jr., and Emmet F.
Low, Jr. October 1955. 20p. diagrs.
(NACA TN 3544)
Comparisons are made between a theory for calcu-
lating stresses about rectangular cutouts in circular
cylinders of semimonocoque construction published
in NACA TN 3200 and previously published NACA
experimental data. The comparisons include
stresses in the stringers and shear stresses in the
center of the shear panels in the neighborhood of the
cutout. The theory takes into account the bending
flexibility of the rings in the structure, and this
factor is found to be important in the calculation of
stresses about cutouts. In general, when the ring
flexibility Is considered, good agreement is exhibited
between the calculated and experimental results.
NACA TN 3550
MEASUREMENTS OF THE EFFECT OF TRAILING-
EDGE THICKNESS ON THE ZERO-LIFT DRAG OF
THIN LOW-ASPECT-RATIO WINGS. John D.
Morrow. November 1955. Ilp. diagram photo.
(NACA TN 3550. Supersedes RM L50F26)
Results of an exploratory free-flight investigation at
zero lift of several rocket-powered drag-research
models having 4-percent-thlck wings of taper ratio
0.423 are presented for a Mach number range of 0.7
to 1.6. Four wings having trailing edges of different
thicknesses were tested. The drag of all the models
was measured and is compared with calculated
NACA TN 3557
A THEORETICAL ANALYSIS OF THE FIELD OF A
RANDOM NOISE SOURCE ABOVE AN INFINITE
PLANE. Peter A. Franken, Massachusetts Insti-
tute of Technology. November 1955. 20p. diagrs.
(NACA TN 3557)
The sound field about a random noise source above a
plane as measured by a receiver with finite band
width is studied theoretically. For simplicity, only
the far field is considered. The special case of a
perfectly reflecting plane is discussed first and the
analysis is then extended to include the case of a
plane of arbitrary impedance.
NACA TN 3567
STUDY OF SCREECHING COMBUSTION IN A
6-INCH SIMULATED AFTERBURNER. Perry L.
Blackshear. Warren D. Rayle, and Leonard K.
Tower. October 1955. 58p. diagrs., photos., tab.
(NACA TN 3567)
The mode of oscillation in a screeching 6-inch-
diameter simulated afterburner is identified through
axial, circumferential, and diametric surveys of
sound amplitude and phase. This mode is found to
be the first Iransverse (slosningi mode in the hot
gases downstream of the flame-holder. The devel-
opment oi a microphone probe suitable for use in
screeching combustors is described. This develop-
ment includes a theoretical and experimental treat-
ment of the attenuation of high-amplitude sound in
NACA TN 3572
AMPLITUDE OF SUPERSONIC DIFFUSER FLOW
PULSATIONS. William H. Sterbentz and Joseph
Davids. October 1955. 23p. diagrs.
(NACA TN 3572. Supersedes RM E52I24)
A theoretical method for evaluating the stability
characteristics and the amplitude and frequency of
pulsation of ram-jet engines without heat addition is
presented. Theory and experiment show that the
pulsation amplitude of a high-mass-flow-ratio
diffuser increases with decreasing mass flow. The
theoretical trends for changes in amplitude, frequen-
cy, and mean pressure recovery with changes in
plenum-chamber volume were experimentally con-
firmed. For perforated, convergent-divergent-type
diffusers, theory and experiment show the existence
of a stability hysteresis loop on the pressure-
recovery, mass-flow-ratio curve.
NACA TN 3573
EFFECT OF EXHAUST-NOZZLE EJECTORS ON
TURBOJET NOISE GENERATION. Warren J. North
and Willard D. Coles. October 1955. 26p. diagrs.,
photo. (NACA TN 3573)
Engine noise levels and jet velocity profiles have
been obtained with several turbojet exhaust-nozzle
ejectors. An insignificant reduction in total sound
power was realized. At subsonic nozzle pressure
ratios, total sound power from exhaust-nozzle
ejectors or bypass exit configurations can be calcu-
lated from primary-jet parameters only.
NACA TN 3575
BURNING VELOCITIES OF VARIOUS PREMIXED
TURBULENT PROPANE FLAMES ON OPEN
BURNERS. Paul Wagner. October 1955. 32p.
diagrs., photos., tab. (NACA TN 3575)
Turbulent burning velocities were measured as a
function of Reynolds number for open propane
flames. Flames of propane and oxygen diluted with
nitrogen, argon, or helium were studied in a variety
of burners up to a maximum pipe Reynolds number
of 26,000. The ratio of turbulent to laminar burning
velocity correlates with the cold-flow Reynolds num-
ber for systems of a given diluent. This ratio also
correlates with a Reynolds number calculated from
values of the turbulent intensity measured at the
RESEARCH ABSTRACTS NO.92
Royal Aircraft Establishment (Gt. Brit.)
LOW SPEED WIND TUNNEL CALIBRATION OF A
Mk. 9A PITOT-STATIC HEAD. J. E. Nethaway.
March 1955. 8p. diagrs., tab. (RAE Tech. Note
This note describes the calibration of a Mark 9A
pitot-static head in the No. 2 11-1/ 2 foot wind
tunnel. The results show the variation of pilot and
static-pressure error coefficients with incidence, at
constant tunnel speed.
Royal Aircraft Establishment (Gt. Brit.)
THE DETERMINATION OF THE RELATIVE REAC-
TIVITIES OF BIFUNCTIONAL MONOMERS. PART
I THE NATURE OF THE PROBLEM. G. M.
Bristow. July 1955. 6p. tab. IRAE Tech. Note
An account is given of the use of functional mono-
mers in the production of cross-linked polymers,
and of some methods by which the relative reactivi-
ties of the two unsaturated groups can be determined.
Royal Aircraft Establishment (GI Brit )
A LOW-PRESSURE MICRO-ANALYTICAL METHOD
FOR DETERMINING OXYGEN, HYDROGEN AND
NITROGEN IN METALS. H. C. Davis and J. A.
Gray. March 1955. 14p. diagrs.. tabs. (RAE
A low-pressure microanalytical method for the
determination of oxygen, hydrogen, and nitrogen in
metals is described. Gases evolved by vacuum-
fusion are analyzed physically. The results obtained
for oxygen and hydrogen are closely reproducible,
but those for nitrogen are less so
RESEARCH ABSTRACTS NO. 92
Royal Aircraft Establishment (Gt. Brit.)
THE THEORETICAL WAVE DRAG OF OPEN NOSE
AXISYMMETRICAL FOREBODIES WITH VARYING
FINENESS RATIO, AREA RATIO AND NOSE ANGLE.
J. H. Willis and D. G. Randail. February 1955.
33p. diagrs., tab. tRAE Tech Note Aero 2360)
Existing results for the wave drag of open-nose
axisymmetrical forebodies are for bodies whose
profiles are straight lines or parabolic arcs. These
results are here extended to a family of profiles
which includes the straight line and the parabolic arc
as special cases. Slender body theory is employed
Royal Aircraft Establishment (Gl. Brit.)
TESTS ON HEAT AND CORROSION RESISTING
COATINGS FOR MAGNESIUM ALLOYS. J. Mackay
and H. G. Cole. June 1955. 9p. tabs. (RAE
Tech. Note Chem. 1254)
In tests on painted magnesium alloy specimens
heated to 2000 C at three-monthly intervals during
an intermittent seawater spray test, the best com-
binations of protective efficiency with hardness and
flexibility were given by an epoxy base scheme and
by a standard stoving scheme to D. T. D. 235.
Other epoxy schemes gave good protection but were
relatively inflexible. A standard air-drying scheme
to D. T.D. 260A gave very good protection. Butyl
titanate paints and a silicone paint gave poor pro-
tection. A zinc chrome sealed anodic treatment
gave better results as a protective pretreatment
than chromate bath III of D. T. D 911A. Butyl titan-
ate paints broke down on pure aluminum exposed to
the same test
EXPERIMENTAL DETERMINATION OF THE EF-
FECTIVE WIDTH OF FLAT PLATES IN THE ELAS-
TIC AND PLASTIC RANGE. (De experimentele
bepaling van de meedragende breedte van vlakke
platen in het elastische en het plastische gebied).
J. F. Besseling. September 1955. 100p. diagrs.,
photos., tab ITrans from Nationaal Luchtvaart-
laboratorium, Amsterdam, S.414, February 19531
Influence of boundary conditions and transverse
stiffeners on the effective width of a plate are dis-
cussed. The testing apparatus, measuring equip-
ment, and lest program are reviewed and test re-
sults are presented for aluminum alloy 24ST.
HEAT-RESISTANT SENTER MATERIALS.
IHochwarmleste Sinterwerkstofle). F. Benesovsky
019551. 6p diagrs photo., tab. (Trans. from
Werkstolle und Korrosion, Jour no. 8 9, Aug -
Sept 1954, p 288-290)
Ceramal materials, produced by powder metallurgy
from a titanium carbide base, are suggested for use
in turbine and rocket drives Other ceramal mate-
rials suggested for nigh temperature use are alloys
with boride. silicide, and oxide bases Titanium
bonrde and zirconium boride show high melting
points, are generally brittle, and have poor thermal
shock properties. It is believed that metal additions
may improve these properties.
NACA Rept 1135
Errata No. 2 on "EQUATIONS, TABLES, AND
CHARTS FOR COMPRESSIBLE FLOW." Ames
Research Staff. 1953.
NACA Rept. 1175
Errata on "EFFECT OF VARIABLE VISCOSITY AND
THERMAL CONDUCTIVITY ON HIGH-SPEED SLIP
FLOW BETWEEN CONCENTRIC CYLINDERS.'
T. C Lin and R. E. Street. 1954
NACA TN 3454
Errata on "EFFECT OF A DISCONTINUITY ON
PARAMETERS WITH APPLICATION TO SHOCK-
INDUCED SEPARATION. Eli Reshotko and
Maurice Tucker. May 1955.
NACA TN 3483
Errata on "AN ANALYSIS OF ACCELERATION,
AIRSPEED, AND GUST-VELOCITY DATA FROM A
FOUR-ENGINE TRANSPORT AIRPLANE IN OPERA-
TIONS ON AN EASTERN UNITED STATES ROUTE."
Thomas L. Coleman and Mary W Fetner.
NACA TN 3514
Errata on "RESPONSE OF HOMOGENEOUS AND
TWO-MATERIAL LAMINATED CYLINDERS TO
SINUSOIDAL ENVIRONMENTAL TEMPERATURE
CHANGE, WITH APPLICATIONS TO HOT-WIRE
ANEMOMETRY AND THERMOCOUPLE PYROME-
TRY." Herman H. Lowell and Norman Patton.
SOME ELEVA ED TEMPERATURE STRUCTURAL
PROBLEMS OF HIGH-SPEED AIRCRAFT.
Richard R. Heldenfels. (Presented to SAE
Golden Anniversary National Aeronautic Meeting,
Los Angeles, California, October 11-15, 1955) 29p.
This paper contains a discussion of the probleiis of
structural design, such as creep, thermal buckling,
thermal stresses, and stiffness reduction, as related
to the aerodynamic heating of high-speed aircraft.
DECLASSIFIED NACA REPORTS
THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL, 10/14/55
NACA RM A7J02
THE HIGH-SPEED AERODYNAMIC EFFECTS OF
MODIFICATIONS TO THE WING AND WING-
FUSELAGE INTERSECTION OF AN AIRPLANE
MODEL WITH THE WING SWEPT BACK 350. Lee
E. Boddy and Charles P. Morrill, Jr. February 18,
1948. 34p. diagrs., photos. (NACA RM A7J02)
Wind-tunnel tests at high subsonic Mach numbers
were conducted on an airplane model having swept-
back wings. Attempts were made to reduce the
interference at the plane of symmetry of the swept-
back wing and thus increase its divergence Mach
number. Also, tests were made with the wing
trailing-edge angle decreased, in an effort to elimi-
nate the reversal of aileron hinge moment and wing
pitching moment suffered by the true-contour wing
at high Mach numbers.
RESEARCH ABSTRACTS NO. 92
NACA RM A7J03
THE AERODYNAMIC EFFECTS OF ROCKETS AND
FUEL TANKS MOUNTED UNDER THE SWEPT-
BACK WING OF AN AIRPLANE MODEL. Lee E.
Boddy and Charles P. Morrill, Jr. April 23, 1948.
19p. diagrs. (NACA RM A7J03)
The effects of externally mounted rockets and fuel
tanks on the aerodynamic characteristics of an air-
plane model with sweptback wings are presented in
this report. Wind-tunnel tests were made at high
subsonic Mach numbers to determine the effect of
the external equipment on the drag, the longitudinal
stability and control, and the lateral control of the
NACA RM A50J26
EXPERIMENTAL DAMPING IN PITCH OF 450 TRI-
ANGULAR WINGS. Murray Tobak. David E. Reese,
Jr., and Benjamin H. Beam. Diccerrmb'r 1, 1950.
63p. diagrs., photo. (NACA RM1 A50J26)
Results are presented of a wind-tunnel investigation
of the variation of the damping-in-pitch parameter
Cm + Cm& with Mach number and axis 1o rotation
position for two triangular wings having leading
edges swept back 450, with and without a body.
Tests were conducted at subsonic speeds over a
Mach number range of 0.23 to 0.94 and at supersonic
speeds from 1.15 to 1.70. The measured damping
coefficients are compared with theoretical results at
both subsonic and supersonic speeds.
NACA RM A53J02
PRELIMINARY RESULTS OF AN INVESTIGATION
OF THE EFFECTS OF SPINNER SHAPE ON THE
CHARACTERISTICS OF AN NACA D-TYPE COWL
BEHIND A THREE-BLADE PROPELLER, IN-
CLUDING THE CHARACTERISTICS OF THE PRO-
PELLER AT NEGATIVE THRUST Robert M
Reynolds. November 1953. 15p diagrs photo
tab. (NACA RM A53J02)
Preliminary results of measurements of the ram-
recovery ratio at the inlet of an NACA D-type cowl
behind an operating propeller in combination with a
1-series and a modified-conical spinner, maximum
efficiency of the propeller with the 1-series spinner
and the spinner-cowling combinations. and the nega-
tive thrust characteristics of the propeller at low
speeds are summarized. Tests were conducted at
Mach numbers from 0.2 to 0.8, for propeller blade
angles from 330 to 630, and for various inlet-
velocity and advance ratios. Negative-thrust char-
acteristics of the propeller were measured at a Mach
number of 0.15 for blade angles from 250 to -200.
All tests were made with the model at an angle of
attack of 00 and at a Reynolds number of 1.0 million
RESEARCH ABSTRACTS NO.92
NACA RM A53J07
THE EFFECTS OF HORIZONTAL-TAIL HEIGHT
AND A PARTIAL-SPAN LEADING-EDGE EXTEN-
SION ON THE STATIC LONGITUDINAL STABILITY
OF A WING-FUSELAGE-TAIL COMBINATION
HAVING A SWEPTBACK WING. Angelo Bandettinm
and Ralph Selan. March 1954. 54p. diagrs.,
photos.. 2 tabs. INACA RM A53J07)
Test results are presented to show the effects of
horizontal-tail height (22- and 8-percent semispan
above wing chord plane extended) on the static longi-
tudinal stability of a model having a wing with 350
sweepback, an aspect ratio -.5, and a taper ratio 0.5
The model was also moduied by a wing-leading-edge
chord extension and tested with the tail in the low
position. Tests were conducted at various Mach
numbers up to 0 92 at a Reynolds number of 2,000.000
and at a Mach number of 0.20 at a Reynolds number
of 11,000,000. The results of airstream surveys in
the region of the tail are also presented.
NACA RM E50J24
DYNAMIC INVESTIGATION OF TURBINE-
PROPELLER ENGINE UNDER ALTITUDE CONDI-
TIONS. Richard P. Krebs, Seymour C. Himmel,
Darnold Blivas, and Harold Shames. December 6,
1950. 55p diagrs., photo. INACA RM E50J241
An altitude-wind-tunnel investigation of the dynamics
of a turbine-propeller engine employing the
frequency -response technique was conducted over a
range of pressure altitudes from 10,000 to 30,000
feet. The dynamic responses generalized for pres-
sure altitude= over the range of Irequencies investi-
gated. The generalized time constants were iound to
be approximately 1.0 second for the engine-propeller
combination, 0.36 second for the propeller alone, and
2.4 seconds for the engine alone. These values were
in good agreement with those predicted Irom steady-
NACA RM E51JII
IGNITION-DELAY CHARACTERISTICS IN MODIFIED
OPEN-CUP APPARATUS OF SEVERAL FUELS
WITH NITRIC ACID OXIDANTS WITHIN TEMPERA-
TURE RANGE 700 TO -1050 F Riley 0. Miller
December 1951. 30p. diagrs., 4 tabs.
(NACA RM E51JIl)
Fluid properties and low-temperature ignition delays
were obtained for approximately 90 fuel-oxidant com-
binations. A red fuming nitric acid containing ap-
proximately 3 percent water and 19 percent nitrogen
tetroxide froze at approximately -870 F and ignited
several low-viscosity luel blends of aromatic amines
in trethylamine at -760 F and lower. With this acid,
the following average ignition delays were obtained
with a blend of 30 percent o-toluidine in triethyla-
Temperature, OF 70 -40 -76 -87 -1051
I Delay, milliseconds 19 24 38 61 1 210j
NACA RM E52J27
PRELIMINARY INVESTIGATION OF A PERFO-
RATED AXIALLY SYMMETRIC NOZZLE FOR
VARYING NOZZLE PRESSURE RATIOS. Eli
Reshotko January 1953. 43p. diagrs photo.,
2 tabs. (NACA RM E52J27T
The performance characteristics of a perforated
axially symmetric contergent -divergent nozzle were
investigated in an attempt to achieve improved
convergent-dnergent nozzle thrust performance at
below design pressure ratios The purpose of the
perforations was to allow inflow of air into the over-
expanded portion of the nozzle, thus advancing sepa-
ration of the flow. The flow through the perfora-
tions was found to advance separation only when Ine
perforations were liberally placed over the entire
divergent portion of the nozzle A local concentra-
tion of perforations caused separation only in the
local region of perforation. The use of low energy
atmospheric bleed air reduced thrust losses by as
much as 50 percent at appreciably overexpanded
operation. For underexpanded flow, air flowing out
through the perforations caused significant thrust
loss. With shrouding to prevent this outbleed,
thrusts 5 to 10 percent less than those of the un-
perforated nozzle were obtained. The use of high
energy bleed was unsatisfactory since the inlet
momentum penalty of the bleed air was in many
cases greater than the additional thrust obtained.
NACA RMI L8129
PRELIMINARY RESULTS OF NACA TRANSONIC
FLIGHTS OF THE XS-I AIRPLANE WITH 10-
PERCENT-THICK WING AND B-PERCENT-THICK
HORIZONTAL TAIL. Hubert M. Drake, Harold R.
Goodman, and Herbert H. Hoover. October 13, 1948.
18p. diagrs.. photos. (NACA RM L8I29)
Contains results of exploratory flights at altitudes of
about 40,000 feet to a maximum Mach number of
1.06. Data are presented showing the longitudinal
trin changes, elevator effectiveness in producing
acceleration, and rudder effectiveness as a function
of Mach number. Data on lateral oscillations are
NACA RM L8J12
HIGH-SPEED WIND-TUNNEL INVESTIGATION OF
A SWEPTBACK WING WITH AN ADDED TRIANGU-
LAR AREA AT THE CENTER. Beverly Z. Henry,
Jr. January 14, 1949. 24p. diagrs., tabs.
INACA RM L8J12)
Results are presented of an investigation m the
Langley 8-loot high-speed tunnel of two sweptback
wings of different plan form. The purpose of the
investigation was to determine the effects of the
addition of a triangular area to the inboard section
of a conventional sweptback wing in order to produce
a wing employing two stages of sweepback. Lift,
drag, and pitching-moment characteristics are pre-
sented to Illustrate these effects for a Mach number
range of 0.40 through 0.935.
NACA RM L8K23
FREE-FLIGHT INVESTIGATION AT TRANSONIC
AND SUPERSONIC SPEEDS OF THE ROLLING EF-
FECTIVENESS OF SEVERAL AILERON CONFIGU-
RATIONS ON A TAPERED WING HAVING 42.70
SWEEPBACK. Carl A. Sandahl. January 11, 1949.
23p. diagrs., photos., tab. (NACA RM L8K23)
An investigation was made of several aileron modi-
fications in conjunction with a tapered, sweptback
wing having circular-arc airfoil sections of rela-
tively large thickness ratio. The modified ailerons
eliminated the reversal of rolling effectiveness at
transonic speeds obtained with the true-contour ai-
lerons at small deflections.
NACA RM L9F07
PRELIMINARY THEORETICAL AND FLIGHT IN-
VESTIGATION OF THE LATERAL OSCILLATION
OF THE X-l AIRPLANE. Hubert M. Drake and
Helen L. Wall. July 19, 1949. 24p. diagrs., photo.,
tab. (NACA RM L9F07)
A small-amplitude, undamped, lateral oscillation has
been encountered in flight tests of the X-1 airplane.
The oscillation occurs in subsonic and supersonic
flight, in maneuvers, and power on and off. The
calculations indicate that a change, in the positive
direction, of the inclination of the principal axis with
respect to the flight path should have a considerable
NACA RM L9G19a
MEASUREMENTS OF AILERON EFFECTIVENESS
OF THE BELL X-l AIRPLANE AT MACH NUMBERS
BETWEEN 0.9 AND 1.06. Hubert M. Drake.
August 4, 1949. 5p. diagrs. (NACA RM L9Gl9a)
Presents results of flight measurements of aileron
effectiveness of the X-1 airplane up to a Mach num-
ber of 0.94. The data indicate a 75 percent loss of
aileron effectiveness between M = 0.82 and
M = 0.94.
NACA RM L50G20
ELEVATOR-STABILIZER EFFECTIVENESS AND
TRIM OF THE X-1 AIRPLANE TO A MACH NUM-
BER OF 1.06. Hubert M. Drake and John R. Carden.
November 1, 1950. 12p. diagrs.
(NACA RM L50G20)
The relative elevator-stabilizer effectiveness of the
X-l has been determined to decrease from a value of
0.25 at a Mach number of 0.78 to a value of 0.05 at a
Mach number of 1.0. At supersonic speeds the ef-
fectiveness increases. The variation between the
trim curves at various stabilizer settings is caused
by the variation in effectiveness and the fact that the
effectiveness is nonlinear at Mael numbers between
0.94 and 0.97. It was found that. with the elevator
fixed at zero, only about 0.50 of stabilizer movement
would be required to trim through the Mach number
range fronl 0.78 to 1.02.
RESEARCH ABSTRACTS NO.92
NACA RM L50J25
A TRANSONIC-WING INVESTIGATION N THE
LANGLEY 8-FOOT HIGH-SPEED TUNNEL AT HIGH
SUBSONIC MACH NUMBERS AND AT A MACH NUM-
BER OF 1.2. WING- FUSELAGE CONFIGURATION
HAVING A WING OF 600 SWEEPBACK, ASPECT
RATIO 4, TAPER RATIO 0.6, AND NACA 65A006
AIRFOIL SECTION. Raymond B. Wood and Frank F.
Fleming. January 24, 1951. 43p. diagrs., photo.
(NACA RM L50J25)
An investigation was conducted in the Langley 8-foot
high-speed tunnel of the aerodynamic characteristics
of a wing swept back 600 at the quarter chord, with
aspect ratio 4, taper ratio 0.6, and an NACA 65A006
airfoil section. The tests were conducted through a
Mach number range from 0.6 to 0.96 and at a Mach
number of 1.2. Data are presented for a wing
fuselage and for a wing with wing-fuselage interfer-
ence. Wake-survey-study results and the measure-
ments of the angle of downwash for a probable tail
location, approximately 38 percent of the wing semi-
span above the wing-chord plane, are included.
NACA RM L51D17
AN INVESTIGATION OF FOUR WINGS OF SQUARE
PLAN FORM AT A MACH NUMBER OF 6.86 IN THE
LANGLEY 11-INCH HYPERSONIC TUNNEL.
Charles H. McLellan, Mitchel H Bertram, and
John A. Moore June 1951 47p diagrs., photos.
(NACA RM L51D17J
The results of tests of four wings at a Mach number
of 6.86 in the Langley 11 -inch hypersonic tunnel are
presented. The wings tested had 4-inch square plan
forms with 5-percent-lhick diamond, half-diamond,
wedge, and half-circular-arc sections. The bound-
ary layer has been found to have a large effect on the
wing pressure distributions Reasonable agreement
was indicated between the aerodynamic coefficients
from experimental pressure data and inviscid theory.
Total drag measurements showed good agreement
with the theory at low angles of attack when the ef-
fects of surface friction were included At the
higher angles of attack, both lift coefficient and drag
coefficient were found to be slightly below the values
predicted by the two-dimensional theory.
NACA RM L51J10
INVESTIGATION BY THE TRANSONIC-BUMP
METHOD OF A 350 SWEPTBACK SEMISPAN MODEL
EQUIPPED WITH A FLAP OPERATED BY A SERIES
OF SERVOVANES LOCATED AHEAD OF AND
GEARED TO THE FLAP William H. Phillips and
Robert F. Thompson December 1951. 39p.
diagrs., photo INACA RM L51J10)
Lift, drag, pitching-moment. rolling-moment, and
yawing-moment data in the Mach number range from
0.6 to 1.0 obtained from wind-tunnel tests of a low-
aspect-ratio sweptoack airfoil model with a servo-
vane control are presented The control utilizes
the drag force and spoiler action of a set of vanes to
deflect a flap-type control Comparison of lift in-
crement and cenler-ol-pressure location is made
-iih previously published data from tests of a con-
ventional flap-lvpe control
RESEARCH ABSTRACTS NO. 92 3
NACA RM L51J30
SUMMARY OF FLUTTER EXPERIENCES AS A
GUIDE TO THE PRELIMINARY DESIGN OF LIFTING
SURFACES ON MISSILES. Dennis J. Martin
November 1951. 16p. diagrs. (NACA RM L51J30)
This report presents a limited review of some ex-
periences in flight testing of missiles and of wing
flutter investigations that may be of interest in mis-
sile design. Several types of flutter which may be of
concern m missile studies are briefly described
Crude criteria are presented for two of the most
common types of flutter to permit a rapid estimate
to be made of the probability of the occurrence of
flutter. Many of the details of the flutter picture
have been omitted, and only the broader elements
have been retained so as to give the designer an
overall view of the subject.
NACA RM L52D01
A STUDY OF THE FLOW OVER A 450 SWEPTBACK
WING-FUSELAGE COMBINATION AT TRANSONIC
MACH NUMBERS. Richard T Whitcomb and
Thomas C. Kelly. June 1952 60p. diagrs ,
photos. (NACA RM L52D01)
Pressure distributions, tuft patterns, and schlieren
surveys have been obtained for a 450 sweptback
wing-fuselage combination in the Langley 8-foot
transonic tunnel at transonic Mach numbers to 1.11
and angles of attack to 200. The results provide an
indication at transonic Mach numbers of the nature
of the formation of shock waves on the wing and
fuselage, wing-fuselage interference, and the de-
velopment of separation and the separation vortex.
NACA RM L52J21a
INVESTIGATIONS AT SUPERSONIC SPEEDS OF THE
BASE PRESSURE ON BODIES OF REVOLUTION
WITH AND WITHOUT SWEPTBACK STABILIZING
FINS. Eugene S. Love and Robert M. O'Donnell.
December 1952. 66p. diagrs., photos.
(NACA RM L52J21a)
Results are presented from an investigation at Mach
numbers of 1.62. 1.93, and 2.41 of the variation
with Reynolds number of the base pressure on bodies
of revolution at zero lift, with and without sweptback
stabilizing fins. Included are the effects of varying
nose and base shapes and cutoff length, the effects
of the presence of sting supports of varying diam-
eter, and the effects of disturbances entering the
wake. The overall Reynolds number range was ap-
proximately from 1 x 106 to 10 x 106.
NACA RM L52J23a
INVESTIGATION OF THE AERODYNAMIC CHARAC-
TERISTICS OF THE NACA RM-10 MISSILE (WITH
FINS) AT A MACH NUMBER OF 1.62 IN THE
LANGLEY 9-INCH SUPERSONIC TUNNEL. Donald
E. Coletti. December 1952. 21p. diagrs.
(NACA RM L52J23a)
An investigation was made of a 0.050-scale model of
the RM-10 missile at a Mach number of 1.62 and a
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8153 288 8
Reynolds number of 2.66 x 106. Measurements were
made of lift, drag, and pitchmg moment over an
angle-of-attack range of *50. The effects of the
ratio of sting-shield diameter to base diameter were
also investigated. Comparisons are made with re-
sults of tests m other facilities at widely different
NACA RM L52K06
PRESSURE DISTRIBUTION AND PRESSURE DRAG
FOR A HEMISPHERICAL NOSE AT MACH NUM-
BERS 2.05. 2.54, AND 3.04. Leo T. Chauvm.
December 1952. 14p. diagrs., photos.
(NACA RM L52K06)
An experimental investigation of the pressure dis-
tributions on a hemispherical nose 3.98 inches in
diameter, mounted on a cylindrical support, has
been made at Mach numbers of 2.05, 2 54, and 3.04
and for Reynolds numbers of 4.44 x 106, 4.57 x 106,
and 4.16 x 106, respectively. The Reynolds number
was based on body diameter and free-stream condi-
tions Pressure-drag coefficients were calculated
and good agreement was obtained between these tests
and other investigations.
NACA RM L53E04
AILERON AND ELEVATOR HINGE MOMENTS OF
THE BELL X-l AIRPLANE MEASURED IN TRAN-
SONIC FLIGHT. Hubert M. Drake and John B
McKay. June 1953. 27p. diagrs.
(NACA RM L53E04)
Hinge moments have been measured on the aileron
and elevator of the Bell X-1 airplane having the 10-
percent-thick wing and 8-percent-thick tail. The
aileron measurements were made by means of
strain gages and pressure distributions while the
elevator measurements were made by means of the
wheel-force strain gages. The elevator hinge-
moment characteristics were determined to a Mach
number of 1.18 and the aderon hinge moments to a
Mach number of 1 13.
NACA RM L53FOB
FLIGHT MEASUREMENTS OF LIFT AND DRAG
FOR THE BELL X-l RESEARCH AIRPLANE HAV-
ING A 10-PERCENT-THICK WING. Edwin J.
Saltzman. September 1953. 37p. diagrs., tab.
(NACA RM L53F08)
Lift and drag results have been obtained from power-
off flight tests of the Bell X-I (10-percent-thick
wing) airplane for Mach numbers 0.68 to 1.01. Com-
parisons of drag are made with 8-percent-thick-wing
flight tests and 10-percent-thick-wing wind-tunnel
S UNIVERSITY OF FLORIDA
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NACA RM L53I09a
LOW-SPEED WIND-TUNNEL INVESTIGATION OF A
JET CONTROL ON A 350 SWEPT WING. John G.
Lowry and Thomas R. Turner. October 1953. 9p.
diagrs. (NACA RM L53I09a)
A low-speed wind-tunnel investigation was made of a
jet control that obtains its effectiveness from both
the jet reaction and from the change in circulation
around the wing due to the jet's acting as a spoiler.
The jet control was investigated as an aileron on a
350 sweptback wing of aspect ratio 4.76. The in-
vestigation was of exploratory nature and was limited
to the case where the jet was supplied with air at
stagnation pressure. The results indicated that such
a jet could be used as an emergency control.
NACA RM L53J01a
WIND-TUNNEL INVESTIGATION OF THE EFFECTS
OF STEADY ROLLING ON THE AERODYNAMIC
LOADING CHARACTERISTICS OF A 450 SWEPT-
BACK WING AT HIGH SUBSONIC SPEEDS. James
W. Wiggins and Richard E. Kuhn. November 1953.
22p. diagrs., photos. (NACA RM L53J01a)
The aerodynamic loading characteristics of a 450
sweptback wing of aspect ratio 4 in combination with
a fuselage during steady roll are presented. The
tests covered Mach numbers of 0.70, 0.85, and 0.91,
and angles of attack up to 130. The effects of fences
at a Mach number of 0.85 and a comparison of
measured and calculated load distribution are in-
NACA RM L53J09a
WIND-TUNNEL INVESTIGATION AT LOW SPEED
OF THE EFFECT OF VARYING THE RATIO OF
BODY DIAMETER TO WING SPAN FROM 0.1 TO 0.8
ON THE AERODYNAMIC CHARACTERISTICS IN
PITCH OF A 450 SWEPTBACK-WING-BODY
COMBINATION. Harold S. Johnson. November
1953. 32p. diagrs., photo., tab.
(NACA RM L53J09a)
Low-speed lift, drag, and pitching-moment data
were obtained for a family of bodies and wing-body
combinations to determine the effect of varying the
ratio of body diameter to wing span from 0.1 to 0.8.
The bodies had ogival noses and cylindrical after-
bodies. The untapered 6-percent-thick wings had
aspect ratios of 3 and 450 of sweepback. The lift
data of the bodies alone and the wing-body combina-
tions are compared with several existing theories.
NACA RM L53J19
AN EXPERIMENTAL AND THEORETICAL INVES-
TIGATION AT HIGH SUBSONIC SPEEDS OF THE
EFFECTS OF HORIZONTAL-TAIL HEIGHT ON THE
AERODYNAMIC CHARACTERISTICS IN SIDESLIP
OF AN UNSWEPT, UNTAPERED TAIL ASSEMBLY.
Harleth G. Wiley and Donald R Riley. December
1953. 71p. alagrs. tab. (NACA RM L53J19)
This paper presents the effects at high subsonic
speeds of horizontal-tall height on the aerodynamic
characteristics in sideslip at 0 angle of attack of an
RESEARCH ABSTRACTS NO.92
unswept, untapered empennage. Configurations in-
vestigated included the fuselage alone, fuselage plus
vertical tail, fuselage plus horizontal tail, and the
fuselage plus vertical tail with the horizontal tail
located at 0, 26, 59, and 100 percent vertical-
surface span. Tests were made at 00 angle of attack
through a sideslip range of -2o to 200 over a Mach
number range of 0.50 to 0.94.
NACA RM L53J29
WIND-TUNNEL INVESTIGATION AT HIGH AND LOW
SUBSONIC MACH NUMBERS OF TWO UNSWEPT
WINGS HAVING NACA 2-006 AND NACA 65A006
AIRFOIL SECTIONS. Stanley F. Racisz.
December 1953. 40p. diagrs., photo., tab.
(NACA RM L53J29)
An investigation has been made of two unswept wings
with aspect ratios of 4 and taper ratios of 0.2. One
wing had airfoil sections designed for high maximum
lift at low speeds (NACA 2-006), and the other wing
had NACA 65A006 airfoil sections Each wing was
mounted on a slender body of revolution. The lift,
drag, and pitching-moment characteristics were
determined at Reynolds numbers from 1 x 106 to
7.5 x 106 for Mach numbers below 0.2 for the wigs
with and without split flaps and for the wings with and
without leading-edge roughness. The character-
istics of the plain wings were also determined for
several values of Reynolds number at Mach numbers
up to about 0.92. Gains obtainable by the use of the
NACA 2-006 airfoil section are evident for Mach
numbers up to 0 65 from the comparisons of the re-
sults for the two wings.
NACA RM L53LOBa
EXPERIMENTAL CONVECTIVE HEAT TRANSFER
TO A 4-INCH AND 6-INCH HEMISPHERE AT MACH
NUMBERS FROM 1.62 TO 3.04. Leo T. Chauvin
and Joseph P. Maloney. February 1954. 18p.
diagrs., photos (NACA RM L53L08a)
Equilibrium temperatures and heat-transfer coeffi-
cients for a hemispherical nose have been measured
for Mach numbers from 1.62 to 3.04. Heat transfer
to the surface of the hemisphere was presented as
Stanton number against Reynolds number for various
surface heating conditions. Heat transfer at the
stagnation point has been measured and correlated
with theory. Transition from a lamiar to a turbu-
lent boundary layer was obtained at Reynolds num-
bers of approximately 1 x 106 corresponding to a
region on the body located between 450 and 600 from
the stagnation point.
NACA RM L53L15
INVESTIGATION OF A PULSE-JET-POWERED
HELICOPTER ROTOR ON THE LANGLEY HELI-
COPTER TEST TOWER. Edward J. Radin and
Paul J. Carpenter February 1954. 23p. diagrs.,
photos. INACA RM L53L15)
A helicopter rotor powered by tip-located pulse-jet
engines has been investigated on the Langley heli-
copter test tower to determine its basic hovering
characteristics as well as the power-off drag and
propulsive characteristics of the engines. The noise
intensity in the vicinity of the pulse-jet engines was
NACA Langley Field. Va.
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