Research abstracts

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Title:
Research abstracts
Physical Description:
93 v. : ; 27 cm.
Language:
English
Creator:
United States -- National Advisory Committee for Aeronautics
Publisher:
National Advisory Committee for Aeronautics
Place of Publication:
Washington, D.C
Publication Date:
Frequency:
irregular
completely irregular

Subjects

Subjects / Keywords:
Aeronautics -- Abstracts -- Periodicals   ( lcsh )
Aeronautics -- Research -- Abstracts -- Periodicals   ( lcsh )
Genre:
serial   ( sobekcm )
federal government publication   ( marcgt )
abstract or summary   ( marcgt )

Notes

Statement of Responsibility:
National Advisory Committee for Aeronautics.
Dates or Sequential Designation:
Abstracts no. 1 (June 15, 1951)-no. 93 (Nov. 30, 1955).

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 001469326
notis - AGY1019
oclc - 01471285
lccn - 86657025
issn - 0499-9274
Classification:
lcc - TL501 .U5895
System ID:
AA00009235:00033

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Monthly list of documents released by the NACA ...
Succeeded by:
Research abstracts and reclassification notice


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Full Text

NI


A method is presented for analyzing the stresses
about a cutout in a circular cylmder of semimono-
coque construction. The method involves the use of
so-called perturbation solutions which are super-
posed on the stress distrabution that would exist in
the structure with no cutout in such a way as to give
the effects of a cutout. The method can be used for
any loading case for which the structure without the
cutout can be analyzed and rs sufficiently versatile to
account for stringer and shear reinforcement about
the cutout.


NACA TN 3213

TRANSONIC FLOW PAST CONE CYLINDERS.
George E. Solomon, Cahforrua Institute of Technology.
September 1954. 56p.diagrs., photos. (NACA
TN 3213)

Expe~rlmental results are presented for transonic
flow past cone-cylinder, axially symmetric bodies.
The drag coelffleent and surface Mach number are
studied as the free-stream Mach number as varied
and, wherever possible, the expe~runental results are
compared wsth theoretical predictions. Interferome-
tric results for several typical flow configurations
are shown and an example of shock-free supersonic-
to-subsonic compression is exp~erlmentally demon-
strated. The theoretical problem of transomc flow
past finite cones Is discussed briefly and an approxi-
mate solution of the axially symmetric transonic
equations, valad for a semi-inflnlte cone, Is presented


NACA TN 3227

APPLICATION OF TWO-DEMENSIONAL VORTEX
THEORY TO THE PREDICTION OF FLOW FIELDS
BEHIND WINGS OF WENG-BODY COMBINATIONS
AT SUBSONIC AND SUPERSONIC SPEEDS. Arthur
Wm. Rogers. September 1954. (II), 1p. diagrs.,
photo., 3 tabs. (NACA TN 3227)

A theoretical method Is evaluated for predlcting flow
fields behind wings of wing-body combinations at
supersonic speeds and slender configurallons at sub-
somec speeds. The method was applied to the calcu-
latlon of downwash at Lall locations behind wings of
triangular-wrng and cylandrical-body combinaltons at
M = 2.0. to illustrate effects of aspect ratio, angle of
attack and incidence, ratso of body radius to wing
semlspan, and angle of Dank on the vortex wake.


WASHINGTON es D C, CITING CODE NUMBER ABOVE EACH TITLE;


NACA TN 320r e-
STRESS ANALYSIS LAR~W\ SEMIMlONO-
COQUE CYLINDERS WITH CUTOUITS BY A PER-
TURBATION LOAD TECHNIQUE. Hlarey G.
McComb, Jr. September 1954. 37p. diagrs., 3tabs.
(NACA TN 3200)


National Advisory Committee FI autics



Research Abstra = oc,
O. 71 ic~iS OCT 7, 19S


CURRENT NACA REPORTS

NACA RM E54G07

INVESTIGATION OF NOISE FIELD AND VELOCITY
PROFILES OF AN A FTERBURNING ENGINE.
Warren J. North, Edmund E. Callaghan and Chester
D. Lanzo. September 1954. 23p. diagrs., photos.
(NACA RM E54G07)

Sound pressure levels and lel velocity profiles are
presented for an engine-afterburner combination
over a range of afterburner fuel-air raltos. At high
fuel-air ratios, a severe low-frequency resonance
producing high noise levels was encountered. A
current fighter aircraft with a different afterburner
confrigurallon yielded considerably lower sound pres-
sure levels and resonance-free operarzon, Indlcating
the importance of acoustic considerations In after-
burning designs.


NACA RM E54G22a

RECOVERY CORRECTIONS FOR BUTT-WELDED,
STRAIGHT-WKIRE THERMOCOUPLES IN HIGH-
VELOCITY, HIGH-TEMPERATURE GAS STREAMS.
Frederick S. Slmmons. September 1954. 19p.
diagrs. (NACA RM E54G22al

Experimental measurements show that a reasonable
correlation among recovery correasions at various
pressures and temperatures for bull-welded
stralght-ware thermocouples Is gaven by an empirical
equation in which the correction is seen to be pro-
portional to the fifth root of the pressure and Inverse-
ly proportional to the fourth root of the temperature.
Resultant probable errors in lemperature measure-
ments are presented and discussed.


NACA TN 3151

EXACT SOLUTIONS OF LAMINAR-BOU~NDARY-
LAYER EQUATIONS WITH CONSTANT PROPERTY
VALUES FOR POROUS WALL WITH VARIABLE
TEMPERATURE. Patrick L. Donoughe and John N.
B. Livingood. September 1954. 42p. diagrs., 2
tabs. (NACA TN 3151)

Solutions were computed for a Prandtl number of 0.7
and a range of coolrng-air flows, and pressure and
wall temperature gradients. For each case, boundary
Layer thicknesses and heal-transfer and frlction co-
efficients were also computed and tabulated. Steeper
temperature profiles for a gaven coolant flow were
obtained by increased wall temperature gradients.
Wall temperature gradnents for zero boundary-layer
tamperature gradients at the wall were Increased by
increased pressure gradient and decreased by in-
creased coolant flow.

'AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1sit H ST., NW,
THE REPORT TITLE AND AUTHOR.

(, 2-f. I 3 72.-
2/ 7











NACA TN 3243

THEORETICAL ANALYSIS OF AN AIRPLANE
ACCELERATION RESTRICTOR CONTROLLED BY
NORMAL ACCELERATION, PITCHING ACCELERA-
TION, AND PITCHIING VELOCITY. Christopher C.
Kraft, Jr. September 1954. 42p. diagrs., 3tabs.
(NACA TN 3243)

An acceleration restrictor which limits the elevator
motion has been analyzed. The system has been
assued to be sensitive to two different control sig-
nals. One signal is proportional to normal and pitch-
ing acceleration and the other is a function of normal
and pitching acceleration and pitching velocity. Sev-
eral values of time lag in the device used to stop the
elevator motion have been studied for two center-of-
gravity positions and for forward speeds up to 1,000
feet per second.







NACA TN 3246

AN EXPERIMENTAL INVESTIGATION OF WHEEL
SPIN-UP DRAG LOADS. Benjamin Milwitzky, Dean
C. Lindquist and Dexter M. Potter. September 1954.
18p. diagrs. (NACA TN 3246. Formerly
RM L53EOBb)

This paper presents some recently obtained informa-
tion on landing-gear applied drag loads and on the
nature of the wheel spin-up phenomenon in landing,
based on a program of tests under controlled condi-
tions in the Langley impact basin. In particular, a
study has been made of the n-,ture and variation of
the coefficient of friction between the tire and the
ruwa during the spin-up process. Also, compari-
sons have been made of the various results obtained
in forward-speed impacts, forward-speed impacts
with reverse wheel rotation, spin-up drop tests, and
forward-speed impacts with wheel prerotation.





NACA TN 3262

STARTING AND OPERATING LIlMITS OF TWO
SUPERSONIC WIND TUNNELS UTILIZING AUXIL-
IARY AI INJECTION DOWNSTREAM OF THE
TEST SECTION. Henry R. Hunczak and Morris D.
Rousso. September 1954. 28p. diagrs., photo.
(NACA TN 3262)

Data are presented for tunnels operating at Mach
numbers 3.85, 3.05, and 2.87 over a range of injector.
to-tunnel mass-flow ratios of 0.5 to 1.35. At Mach
number 3.85, the starting pressure ratio of 9.8 with-
out injectors but with a fixed second throat was re-
duced to 4.68 with injectors operating at an injector-
to-tunnel mass-flow ratio of 1.27. The running
pressure ratio was lowered from 8.3 to 4.5. Corre-
sponding reductions at Mach number 3.05 were from
4.5 to 2.71 for starting and from 4.5 to 2.37 for run-
ning at a mass-flow ratio of 0.9.


NACA
RESEARCH ABSTRACTS NO.71


NACA TN 3285

SECTION CHARACTERISTICS OF AN NACA 0006
AIRFOIL WITH AREA SUCTION NEAR THE LEAD-
ING EDGE. James A. Weiberg and Robert E.
Dannenberg. September 1954. 47p. diagrs., photos.,
3 tabs. (NACA TN 3285)

Results are presented of an investigation of a two-
dimensional NACA 0006 airfoil with area suction near
the leading edge. The maximum lift of the wing wlth-
out suction was increased 43 percent for a section
flow coefficient of 0.0010 at a free-stream, velocity
of 162 feet per second.



NACA TN 3287

HEAT TRANSFER FROM A HEMISPHERE-
CYLINDER EQUIPPED WITH FLOW-SEPARATION
SPIKES. Jackson R. Stalder and Helmer V. Nielsen.
September 1954. 29p. diagrs., photos. (NACA
TN 3287)

Average heat-transfer, temperature-recovery, and
pressure-distribution measurements were obtained
over the hemispherical nose of a body of revolution
both with and without flow-separation spikes. The
tests, conducted in a range of Reynolds numbers from
1.55 to 9.85 x 105 and from Mach number 0.12 to
5.04, indicated that in supersonic flow the addition of
spikes approximately doubles the rate of heat trans-
fer, regardless of spike length.



NACA TN 3300

INVESTIGATION OF LIFT, DRAG, AND PITCHING
MOMENT OF A 600 DELTA-WING-BODY COM-
BINATION (AGARD CALIBRATIION MODEL B) IN
THE LANGLEY 9-114CH SUPERSONIC TUNNEL.
August F. Bromm, Jr. September 1954. 18p.
diagrs., photos. (NACA TN 3300)

Results are presented from tests of the AGARD
Calibration Model B in the Langley 9-inch supersomec
tunnel. Measurements were made of the lift, drag,
and pitching moment at Mach numbers of 1.62, 1.94,
and 2.41 and at a Reynolds number, based on body
length, of approximately 3.0 x 106. The zero-lift
drag data compared favorably with available data
and were in the proper sequence for the effects of
Reynolds number.




BRITI1SH REPORTS




N-32662X

Nat. Gas Turbine Establishment (Gt. Brit.)
THE REFLECTION AND TRASMISSION OF PRES-
SURE DISTURBANCES AT AREA CHANGES IN
INFINITE DUCTS. M. V. Nesbitt. February 1954.
103p. diag~rs., tab. (NGTE: Memo.Mi. 198)






NACA
RESEARCH ABSTRACTS NO.71


In most problems Involving the propagation of pres-
sure disturoances, refleenion, and transmission of
the disturbance at area changes must be considered*
In this paper dala are presented describing the re-
flection and transmission of pressure disturbances at
abrupt area changes Inside infinite ducts and at the
ends ofI semialnfinlte ducts. The presentation is such
that only an Initial flow condition, the ratio of the
downst reamr area to the upstream area, the pulse
size and its direction of travel need be defined to
obtain the magnitude of the reflected and transmitted
pulses. Although the results are expressed in terms
tey maynbe rea ly cso~n ere to trs owf r ssre
or velocity. The data refers only to Mlach numbers
Less than 1.





N-32686'

Royal Aircraft Establishment (Gt. Brit.)
A HOT-WIRE ANrEMOMETER FOR THE MEASURE-
MENT OF SMALL AIR FLOWS. P. A. Stickles and
T. R.H. SIzer. March 1954. 16p. diagrs., photos.
(RAE Tech. Note EL. 50)

The principle underlying the operation of the hot-wire
anemometer and its field of application in the meas-
uremient of small air flow are described. Full elec-
Inlcal and constructional details of the instrument are
gaven together with details of a cone-valve for con-
trolling the back-pressure and a low-pressure
chamber for simulating high altitude conditions. The
sources of error are discussed and an indication is
gaven of the Ilmits of accuracy to be expected,







N-32694 "

Royal Aircraft Establishment (Gt. Brit.)
APPROXIMATE THEORETICAL CALCULATIONS
OF THE E FFECT OF CYLINDRICAL TAILPIPES ON
THE THRUST OF A ROCKET MOTOR. D. M.
Clemmow. March 1954. 57p. diagram 16 tabs.
(RAE RPD 18)
Assuming one-dan~ensional flow (free from oblique
shocks) calcularsons have been made of the effect of
nonconducting cylindrical tailpipes on the thrust
delivered by rocket motors. Both supersonic and
subsonic pipes of a range of lengths and diameters
have been examined for a number of different cham-
ber pressures, and for each case optimum area
expansion ratios for motors with pipes have been
determined. With supersonic pipes the thrust loss
increases to a maximum value (of nearly 20 percent
for the worst case investigated) which occurs when
the pipe Is long enough for the flow to become sonic
at its downstream end. With subsonic pipes the
existence of irreversible flow upstream of the nozzle
throat necessitates separate examination of different
cases.


N-32695*

Royal Aireraft Establishment (Gt. Brit.)
STATIC AND DYNAMIC RESPONSE OF A DESIGN OF
DIFFEREGNTIAL PRESSURE YAWMETER AT SUPER-
SONIC SPEEDS. L. J. Beecham and S. J. Collins.
February 1954. 56p. diagrs., photos. (RAE GW 19)

Static calibration curves are provided for incidences
up to 30oat speeds from M = 1.3 to 1.0, and the in-
strument is shown to resolve accurately when rolled
out of the free stream incidence plane. Relations are
developed from which a close approximation to the
Mach number, incidence and roll angle may be ob-
tained, without recourse to calibration curves, in
terms of the differential pressures across each pair
of holes and the pitot pressure measured at an axial
hole; the free stream static pressure requires to be
determined independently. Dynamic behavior of the
instrument and associated pressure pick-ups is ex-
amined, and a design developed for which the acoustic
natural frequency is high compared with that likely
to be encountered in flight.


N-32697*

Nat. Gas Turbine Establishmnent (Gt. Brit.)
THE DESIGN AND TESTING OF A FULL SCALE
VORTEX COMBUSTION CHAMdBER FOR RESIDUAL
OIL BURNIN. A. P. Johnstone. March 1954.
65p. diagrs., photos., 5 tabs. (NGTE Memo.MI.203)

This memorandum describes the design, develop-
ment, and performance of a large vortex chamber
burning 6,000 seconds residual fuel oil. The cham-
her ~was designed to pass 38 lb/sec of air at 4
atmospheres and 100 percent overload tests were
attempted. The chamber was fully cooled by dilution
air for a total loss of 2.0 lb/sq in, at 4 atmospheres.
The! system had three tangential entries and operated
on one, two, or three entries with uniform sym-
metrical exhaust temperature profiles. Combustion
efficiencies were high, 97 to 99 percent, falling off
more rapidly below about 50 percent of the design
mixture strength.




N-32701f

Ministry of Supply (Gt. Brit.)
AN ANALYSIS OF N. A. C.A. HELICOPTER
REPORTS. R. N. Liptrot. Mlay 1954. 61p. diagrs.,
5 tabs. (MOS 8 & TM 7/54)

Theory is compared with flight and model tests, in
order to obtain empiric correcting factors which will
enable reliable performance estimates to be made
for new helicopter designs. A survey of general
theory is followed by an analysis of certain American
reports. Correcting factors for effective blade drag
for tip speed ratio, compressibility and stalling of
the retreating blade are derived. A method of cal-
culating the retreating blade tip angle of attack for
twisted blades is presented. The work was carried
out by the author under Ministry of Supply sponsor-
ship.










N-32723*

Royal Aircraft Establishment (Gt. Brit. )
THE: TENSILE PROPERTIES OF D. T. D. 546 AND
D. T.D. 687 AFTER HEATIG AT ELEVATED
TEMPERATURES. M~ay 1954. 5ip. diagrs. (RAE
Tech. Note Met. 197)

The tensile properties of D. T. D. 546 and D. T. D. 687
after heating at various temperatures up to about
3000 C for periods ranging from 5 seconds to 1000
hours have been reviewed and plotted in convenient
curves. At temperatures of about 1250 C and above,
the aluminum-sine-magnesium alloy D. T. D.687
loses strength to a very much greater extent than the
D. T. D.546, thus losing the higher strength advan-
tage it possesses at room temperature.







N-32725*

Royal Aircraft Establishment (Gt. Brit. )
TENSILE AND COMPRESSIVE TESTS ON TITANIUM
STRIP AT ELEVATED TEMPERATURES. D. C.
Hayward. March 1954. 34p. diagrs., photos.,
12 tabs. (RAE Tech.Note Met. 194)

The compressive testing of hot rolled commercial
purity titanium strip was carried out in a fixture
using steel balls to prevent lateral buckling of the
test piece. Tests in tension and compression at
temperatures ranging from room temperature to
4000 C were made to provide data for tangent moduli
graphs. Longitudinal and transverse tensile strength
values were almost equal but under compression
there was marked anisotropy. Strength was well
maintained at the higher temperatures. Differences
in stiffness were greatest at room temperature but
were appreciably reduced at 4000 C.





N-32726*

Royal Aircraft Establishment (Gt. Brit.)
DESIGN AND DEVELOPMENT OF THE R. A. E.
DUMMY OF THE STANDARD AIRMAN. G. Lovell.
May 1954. 27p. diagrs., photos., 4 tabs. (RAE
Tech. Note MVech. Eng. 176)

A dummy of the standard airman has been developed
at the R. A. E. for use in dynamic and static tests. It
has been designed to the dimensions of the average
airman and the limbs and the complete man have ap-
proximately the correct weights and centers of gravity.
Its all up weight is normally 166 pounds but it can be
varied from this figure if desired. The dummy has a
basic metal structure, covered with foam rubber and
articulated at the joints. While it is normally fitted
with a simple canvas suit, it can be dressed in stand-
ard aircrew clothing, equipment, and headgear.
There are cavities for instruments in the head and in
the upper and lower trunk. A box containing three
mechanical accelerometers mutually at right angles
can be fitted in each of these cavities.


NACA
RESEARCH ABSTRACTS NO. 71


N-32727 *

Royal Aircraft Establishment (Gt. Brit.)
TIHE: STABILITY OF THE: RUNGE-K(UTTA METHO
OF SOLUTION OF LINEAR DIFFERENTIAL
EQUATIONS. R. H. Merson. June 1954. 12p.
diagrs. (RAE: Tech. Note GWT 320)

The solution of a linear differential equation with
constant coefficients as given by the Hunge-Kutta
method is obtained and the stability is studied. It is
shown that the Runge-Kutta solution will be stable
when the true solution is stable if the interval of
integration is less than about 2.8 times the time con-
stant of every subsidence and less than about 2.8
times the reciprocal of the undamped frequency of
every oscillatory mode.



N-32729X

Royal Aircraft Establishment (Gt. Brit.)f
CALCULATION OF CHORD AND SPANWIS
LOADINGS ON THREE THIN WINGS. L. Klner.
April 1954. 13p. diagrs., 4 tabs. (RAE Tech. Note
Aero 2306)

The pressure distribution over three thin wings has
been calculated and compared with results obtained
by Redshaw using his "three-dimensional potential
analyzer'. The agreement is good.



N-32730 X

Royal Aircraft Establishment (Gt. Brit.)
FLEXIBLE EDGE ATTACHMENTS FOR WIND-
SCREENS. F. M. Clowes and L. W. Lord. May
1954. 13p. diagrs., 3 tabs. (RAE Tech. Note
Chem. 1228)

A flexible edge attachment consisting of glass cloth
and polyvinylbutyral has been developed for aircraft
windscreens. While the attachment enables the wind-
screen to be bolted securely to the frame, it will ac-
commodate a difference of thermal expansion between
the frame and the windscreen. The joint strength
decreases with rise in temperature so that the
method is best suited to normal and low temperature
conditions.




MISCELLANEOUS




N-3074D

INDEX OF NACA TECHNICAL PUBLICATIONS,
JUNE, 1953 MIAY, 1954. 1954. viii, 212p.
(NACA)

There is a classified listing of subject categories;
a chronological listing of NACA publications under
each subject category; an alphabelseal index to the
subject categories; and an author index.






NACA
RESEARCH ABSTRACTS NO.71


NACA Rept. 1135

Errata No. 1 on "EQUATIONS, TABLES, AND
CHARTS FOR COMPRESSIBLE FLOWK. Ames
Research Staff. 1953.




NACA TN 3100

Errata No. 1 on aAN INVESTIGATION OF LAMEL-
LAR STR UCTURES AND MINOR PHASES IN
ELEVEN COBALT-BASE ALLOYS BEFORE AND
AFTER HEAT TREATMENT. D J. W. Weeton and
R. A. Signorelli. March 1954.



NACA TN 3118

Errata No. 2 on DESIGN DATA FOR MULTIPOST-
STIFFENED WINGS IN BENDING. Roger A.
Anderson, Aldie E. Johnson, Jr. and Th~omas W.
Wilder, III. January 1954.





UNPUBLISHED PAPERS





N-18639*

ON THE SUITABILITY OF E`LECT`ROLYTICALLY
POLISHED SPECIMENS FOR THE M/ETALLO-
GRAPHIC INVESTIGATION OF ALUMINUM ~AND
ITS AL LDYS. (Ujber die Eignung elektrolytische
polierter Proben fitr die metallographische
Untersuchung von Aluminum und seinen Legierungen).
H. Rohrig and W. Schneider. Augst 1954. 10p.
diagr., photos. (Trans. from Alluminu, v.23, no., ,
June,1941, p. 281-284)

Known methods for electrolytic polishing of aluminum
parts are reviewed and discussed. It is concluded
that the applicability of electrolytic polishing methods
to the preparation of metallographic specimens in
the field of aluminum is comparatively limited.




DECLASSIFIED NACA REPORTS




THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONVFID'ENTIA,
9/10/54


NACA RM A50HO3

PRELIMINARY FLIGHT INVESTIGATION OF THE
~WINVG-DROPPING TENDENCY AND LATERAL-
CONTROL CHARACTERISTICS OF A 350 SWEPT-
WING AIRPLANE AT TRANSONIC MACH NUMBERS.
George A. Rathert, Jr., L. Stewart Rolls, Lee
Winograd and George E. Cooper. September 11.
1950. 14p. diagrs., photo., tab. (NACA RM A50HO3)
(Declassified from Confidential, 9/10/54)

Results are presented from a preliminary flight in-
vestigation on ~an F-86Aairplane of the lateral-
control characteristics and the wing-dropping
tendency encountered at high Mach numbers. The
wing-dropping tendency was found to result from a
combination of three factors: a small initial direc-
tional asymmetry, an abrupt increase in positive
dthedral effect, and a reduction in the! lateral-control
effectiveness.






NACA RM A50JO9a

PRELIMINARY FLIGT INVESTIGATION OF TH
DYNAMIC LONIGITUD~INAL-SjTABILITY CHARAC-
TERISTICS OF A 35o SWEPT-WING AIRPLANE.
Williamn C. Tr$ipett and Rudolph D. Van Dyke, Jr.
December 11, 1950. 26p. diagrs., photo., tab.
(NACA RM A50JOga) (Declassified from Confidential,
9/10/54)
Flight tests were conducted on a 35o swept-wing air-
plane to determine the dynamic longstudmal-stabillity
characteristics. Period and damping of oscillatory
responses were measured over a Mach number range
of 0.60 to 1.04. Also presented as functions of Mach
number are the static stability paramaeter Cm, the
factor Cmg + Omo, and the number of cycles re-
quired for the oscilation to damnp to 1/10 amplitude.
\Sharp variations in damping were noted between Mach
nubers of 0.88 and 0.96 in addition to a decrease
above Mach number 0.96.





NACA RM A51B28

FUGHT MEASUREMENT OF THE WING-
DROPPING TENDENCY OF A STRAIG;HT-WING JEZT
AIRPLANE AT HIGH SUBSONIC MACH NUMBERS.
Seth B. Anderson, Edward A. Ernst and Rudolph D.
Van Dyke, Jr. April 24, 1951. 16ip. diagrs., photo.,
tah. (NACA RM A51B28) (Declashified from
confidential, 9)k0/54)

Flight tests conducted on a straight-wing fighter-type
jet airplane showed that the wing-dropping tendency
encountered at high subsonic Mach numbers was duei
primarily to a progressive reduction in aileron ef-
fectiveness and an increase in effective dihedral
whrlich made thre lateral trim particularly sensitive
to small changes in sidealip angle.


RM A5 1H20
RM L51111









NACA RM A51C28

A COMPARISON OF THE MEASURED AND PRE-
DICTED LATERAL OSCILLATORY CHARACTER-
ISTICS OF A 350 SWEPT-WING FIGHTER AIR-
PLANE. Walter E. McNeill and George? E. Cooper.
Julyl951. 21p.diagrs.,3tabs. (NACARMA51C28)
(Deleassified from Confidential, 9/10/54)
Flight measurements of the lateral oscillatory char-
acteristics of a 350 swept-wing fighter airplane were
obtained at altitudes of 10,000 and 35,000 feet for an
over-all Mach number range from 0.41 to 1.04.
Period, time to damp to half amplitude, and ratio of
angle-of-bank amplitude to angle-of-sidealip ampli-
tude, 1~I l /p I, are presented as functions of Miach
number for each test altitude, and are compared w~ith
values computed from wind-tunnel data and estimated
stability derivatives. In general, agreement between
the measured and predicted lateral oscillatory char-
acteristics of the test airplane was satisfactory.





NACA RM A51E03

SUBSONIC M(ACH AND REYNOLDS NUMBER
EFFECTS ON THE SURFACE PRESSURES, GAP
FLOW, PRESSURE RECOVERY, AND) DRAG OF A
NONROTATING NACA 1-SERIES E-TYPE COWLING
AT AN ANGLE OF ATTACK OF 00. Robert M*
Reynolds and Robert I. Sammonds. July 1951. 73p.
diagrs., photo., 3 tabs. (NACA RM A51EO3)
(Declassified from Confidential, 9/10/54)

Measurements of pressure distributions on the ex-
ternal, inner lip, spinner, and propeller-blade-shank
fairing surfaces are presented for an NACA 1-51-117
(E-type) cowling with an NACA 1-41.413-042.86 spin-
ner suitable for a turbine-propeller installation.
Ram-recovery ratios, external drag, and gap leak
flow are also included. The tests were conducted
with the model at an angle of attack of Go and with
the cowling not rotating. The flow characteristics^
are presented for a range of Mach numbers from.
0.23 to 0.88 at a Reynolds nuber of 1.8 million and
for a range of Reynolds numbers from 1.8 to 8.1 mil-
lion at 0.23 Mach number. Inlet-velocity ratio was
varied between 0.06 and 0.78.



NACA RM A51E14

MEASUREMENTS IN FLIGHT OF THE LONGITUDI-
NAL CHARACTERISTICS OF TWO JET AIRCRAFT,
ONE WITH A DIVING TENDENCY AND THE OTHER
WIITH A CLIMBING TENDENCY AT HIGH MACH
NUMBERS. Seth B. Anderson. October 1951. 18p.
diagrs., photos., 2 tabs. (NACA RM A51E14)
(D~eclassified from. Confidential, 9/10/54)
Flight tests conducted on two jet airplanes of gener-
ally similar configuration, one identified with a div-
mng tendency and the other with a climbing tendency
at high Mach numbers, showed that the difference in
longitudinal control appears to be governed by the
balance between two opposing moments; a diving
tendency caused chiefly by an increase in angle of
attack of the horizontal tail surfaces, and a climbing
tendency due to the pitching moment of the wing.


NACA
RESEARCH ABSTRACTS NO.71


NACA RM A51E24

AN EXPERIMENTAL INVESTIGATION AT SUBSONIC
SPEEDS OF A SCOOP-TYPE AIR-INDUCTION
SYSTEM FOR A SUPERSONIC AIRPLANE. Curt A.
Holshauser. July1951. 45p.diagrs.,photos. (NACA
RM A51E24) (Declassified from Confidential,
9/10/54)

Measurements of ram-recovery ratio, static pressure;
and boundary layer are presented at a Mach number
of 0.17 for a large range of mass-flow ratios, angles
of attack, and angles of sideslip for a scoop-type in-
take located on top of the fuselage. At Go angles of
attack and sideslip the ram-recovery ratio measured
at the minimum-area station was 0.08 at a mass-
flow ratio of 1.0. Above a mass-flow ratio of 1.2, the
ram-recovery ratio decreased rapidly. The variation
of ram-recovery ratio with angle of attack wras small
compared with the variation of ram-recovery ratio
with angle of sideslip.





NACA RM A51Fl2a

INVESTIGATION OF A TRIANGULAR WING IN CON-
JUNCTION WITHI A FUSELAGE AND HORIZONTAL
TAIL TO DETERMINE DOWNWASH AND LONGITUZ-
DINAL STABILITY CHARACTERISTICS TRAN-
SONICBUMP METHOD. Edwin C. Allen. August
1951. 22p. diagrs., photos. (NACA RM A51F12a)
(Declassified from Confidential, 9/10/54)

This report presents effective downwash at the tail
position and the static longitudinal-stability charac-
teristics of a semispan model having a thin triangular
wing of aspect ratio 2, a slender fuselage, and a thin,
unswept horizontal tail. The range of Mach numbers
was from 0.40 to 1.10 with a corresponding Reynolds
number range of 1.0 to 1.9 million. The effects of
vertical position of the horizontal tail at one longitu-
dinal station behind the wing were investigated.






NACA RM A51G27

LONGITUDINAL FREQUESNCY-RESPONSE CHARAC-
TERISTICS OF A 350 SWEPT-WING AIRPLANE AS
DETERMINED FROM FLIGHT MEASUREMENTS,
INCLUDING A METHOD FOR THE EVALUATION OF
TRANSFER FUNCTIONS. William C. Triplett and G.
Allan Smith. September 1951. 4l5p. diagrs., photo.
(NACA RM A51G27) (Declassified from Confidential,
9/10/54)

Data obtained from dynamics flight measurements are
used to compute the longitudinal frequency-response
characteristics of a 350 swept-~wing airplane through
the Mach number range of 0.50 to 1.05. Also pre-
sented is a general graphical method for the deter-
mination of nuerical coefficients of the~ transfer
functions from frequency-response data. The varla-
tions in these coefficients with changes in Mach nu-
ber and altitude are shown.






NACA
RESEARCH ABSTRACTS NO. 71



NACA RM A51HO9

STABILITY AND CONTROL MEASURE ENTS OB-
TAINED DURING USAF-NACA COOPERATIVE
FLIGHT-TEST PROGRAM ON THE X-4 AIRPLANE
(USAF NO. 46-677). Melvm Sadoff, Herman O.
Ankenbruck and Whilliam O'Hare. October 1951. 38p.
diagrs., photos., tab. (NACA RM A51HO9) (Declas-
saf led frojm Conf idential, 9 10 541

Results obtained during the Air Force testing of the
Northrop X-4 airplane are presented. Information is
included on the stalling characterlstics, the static and
dynamic longitudinal- and lateral-directional stability
characterist Ics and the lateralI-cont rol characterrs-
tics.






NACA RM A51H20a

COMPARISON OFDRAG, PRESSURE RECOVERY,
AND SURFACE PRESSURE OF A SCOOP-TYPE IN-
LET AND AN NACA SUBMERGED INLET AT TRAN-
SONIC SPEEDS. Joseph L. Frank and Robert A.
Taylor. December1951. 63p. diagrs., photos.
(NACA RM A51H20a) (Declassified from Confidential,
9 10 54)

Comparative data were obtained for a scoop-type in-
let and an NACA submerged inlet at transonic speeds.
The submerged inlet effected the higher ram re-
covery at mass-flow ratios below about 0.50. At
maximum mass-flow ratio (0.92) and Oo angle of
attack, the ram recovery of the two inlets was about
equal. In general, the scoop-type inlet caused the
greater external drag.








NACA RM A51ll2

A FLIGHT EVALUATION OF THE LONGITUDINA L
STABILITY CHARACTERISTICS ASSOCIATED WITH
THE PITCH-UP OF A SWEPT-WING AIRPLANE IN
MANEUVERING FLIGHT AT TRANSONIC SPEEDS.
Seth B. Anderson and Richard S. Bray. November
1951. 33p. diagrs., photo., tab. (NACA RM A51112)
(Declassified from Confidential, 9 10 54)

Flight measurement s on a swept -wing Jet aircraft
showed that the pitch-up encountered in a wind-up
turn at transonic Mach numbers was due principally
to an unstable break in the wing pitching moment
associated with flow separation near the wing tip.
The pitch-up encountered in slowing down in a dlrve-
recovery maneuver was due chiefly to a reduction in
wing-fuselage stability. An increase in down load
for the horizontal tall was indicated with Increase in
Mpch number for normal force-coefficient values in
excess of approximately 0.2.


NACA RM A51Jll

THE EFFECTS AT TRANSONIC SPEEDS OF THICK-
ENING THE TRAILING EDGE OF A WING WITH A
4-PERCENT-THICK CIRCULAIR-ARC AIRFOIL.
Joseph W. Cleary and George L. Stevens. December
1951. 43p. diagrs., photo. (NACA RM A51711) (De-
classitied from Confidential, 9 10/54)

Effects of systematic \arlatron of trailing-edge thick-
ness of a symmetrical circular-are airfoil on lift,
drag, pitching moment, base pressure, and wake
pressure fluctuations were investigated for the tran-
sonic Mach number range by the wind-tunnel bump
technique. Results show that fortrailing-edge thick-
ness of 0.3 of the airfodl thickness beneficial gains in
lift-drag ratios can be expected at subsonic Mach
numbers witlh no measurable increase in minimum
drag. Higher lift-curve slopes were observed in the
transonic: Mach number range for blunt-trailing-edge
airfoils as compared with circular-are airfoils.




NACA RM A51Jl8

THE EFFECTIVE ENESS OF WING VORTEX GENER-
ATORS IN IMPROVING THE MANEUVERING CHAR-
ACTERISTICS OF A SWEPT-WING AIRPLANE AT
TRANSONIC SPEEDS. Norman M. McFadden, George
A. Rathert, Jr. and Richard 5. Bray. February 1952.
45p. photos., diagrs., tab. (NACA RM A51J18)
(Declassafzed from Confidential, 9/10/54)

The effects of wing vortex generators, multiple
bounda ry-layer fences, and extension of the outer two
segments of the wing leading-edte slats on the aero-
dynamic characteristics of a 35 swept-wing fighter
were measured in light tests at transonic speeds and
high altaludes. Slgnifacant improvements were ob-
tained In the pitch-up and wilng-dropping-tendency
characteristics with certain arrangements of vortex
generators.




NACA RM A51 JI9a

TH E FFFECT OF ENTRANCE MACH NUMBER AND
LIP SHAPE ON THE SUBSONIC CHA~RACTERISTICS
OF A SCOOP-TYPE AIR-INDUCTION SYSTEM FOR
A SUPERSONIC AIRPLANE. Curt A. Holzhauser.
January 1952. 39p. diagrs., photos., tab. (NACA
RM A51J19a) (Declassified from Confidential,
9 10 54)

This report presents the ram-recovery ratios and
static-pressure distributions of a scoop-type intake
with a rounded lip and with a sharp lip. The en-
trance Mach numbers ranged from 0l to choking for
free-stream Mach numbers of 0.08 to 0.33. The
wake drag coefficients of these two types of installa-
tions are compared. The ram-recovery ratio of the
intake with a rounded lip was greater tha with a sharp
lip at the higher mass-flow ratios. When the air flow
was separated in the duct, the ram-recovery ratio de-
creased with increasing entrance Mach numbers.






NACA
RESEARCH ABSTRACTS NO.71


NACA RIM E51CO2

CARBON DEPOSITION OF SEVERAL SPECIAL
TURBOJET-ENGINE FUELS. Jerrold D. Wear and
James W. Useller. April 10l, 1951. 15p. photos.,
diagr., tab. (NACA RM E51CO2) (Declassified from
Confidential, 9/10/54)
Investigations were conducted to determine the car-
bon forming characteristics of MIL-F-5624 and
MIL-F-5161 type fuels in a single J33 combustor and
of a MIL-F-5161 fuel in a J35 full-scale engine. The
carbon deposition of the fuels investigated in the
single combustor could be estimated from a previ-
ously established empirical correlation with volu-
metric average boiling temperature and hydrogen-
carbon ratio. The results indicated thatMIL-F-5161
type fuels formed more carbon in the single comb-
bustorr and the full-scale turbojet engine than most
MIL-F-5624 type fuels, and may result in marginal
operation in several turbojet engines.



NACA RM E51C14

EFFECT OF FUEL VOLATILITY ON PERFORM-
ANCE OF TAIL-PIPE BURNER. Zelmar Barson
and Arthur F. Sargent, Jr. April 30, 1951. 18p.
diagrs., tab. (NACA RM E51C14) (Declassified
from Confidential, 9/10/54)
Fuels having Reid vapor pressures of 6.3 and 1.0
pounds per square inch were investigated in a tail-
pipe burner on an axial-flow-type turbojet engine ata
simulated flight Mach number of 0.6 and altitudes
from 20,000 to 45,000 feet. With the burner con-
figuration used in this investigation, having a mixing
length of only 8 inches between the fuel manifold and
the flame holder, the low-vapor-pressure fuel gave
lower combustion efficiency at a given tail-pipe fuel-
air ratio. Because the exhaust-nozzle area was
fixed, the lower efficiency resulted in lower thrust
and higher specific fuel consumption. The maximum
altitude at which the burner would operate was prac-
tically unaffectedby the change in fuel volatility.



NACA RM E51D16

INVESTIGATION OF MECHANISMS OF BLADE
FAILURE OF FORGED HASTELLOY B AND CAST
SATELLITE 21 TURBINE BLADES IN TURBOJET
ENGINE. C. Yaker, C. F. Robards and F. B.
Garrett. August 1951. 41p. diagrs., photos., 2 tabs.
(NACA RM E51D16) (Declassified from Confidential,
9/10/54)
An investigation was conducted to study the mecha-
nismls of blade failure of forcedsastelloy Band cast
Stellite 21. The blades were mountedin a 16-25-6
alloy rotor and subjected to 20l-minute cycles con-
sisting of 15 minutes at rated speed and approximate-
ly 5 minutes at idle. The first failures of the
Hastelloy B and Stellite 21 blades were probably the
result of excessive vibratory stresses and occurred
after 14.25 and 16.75 hours, respectively. After
28.75 hours of operation, all but 3 of the original 25
Hastelloy B blades had either failed or contained
stress-rupture-type cracks and four of the original
2a Stellite 21 blades contained stress-rupture-type
cracks.


NACA RM E50H16a

INVESTIGATION OF IGNITION CHARACTERISTICS
OF AN-F-32 AND TWO AN-F-58a FUELS IN
SINGLE CAN-TYPE TURBOJET COMBUSTOR.
Warren D. Rayle and Howard W. Douglass.
October 13, 1950. 25p. photos., diagrs., 2 tabs.
(NACA RM E50H16a) (Declassified from Confidential,
9/10/54) '

Ignition characteristics of AN-F-32 and two
AN-F-58a fuels were studied in a single can-type
turbojet combustor under air-flow conditions repre-
senting engine speeds of 1600, 2504), and 4000 rpm,
altitudes from sea level to 30,000 feet, ambient
temperatures at sea level from 900 to -360 F, and
flight Mach numbers of 0 and 0.6. Critical fuel-flow
rates for ignition increased with increase in preigni-
tion engine speed, with increase in altitude, or with
decrease in sea-level ambient temperature. This
flow rate appears to increase in a direct relation to
decrease in fuel volatility as indicated by the 10-
percent-evaporated temperature.






NACA RM E51BO2

COMBUSTION PROPERTIES OF ALUMINUM AS
RAM-JET FUEL. J. Robert Branstetter, Albert M.
Lord and Melvin Gerstein. March 28, 1951. 37p.
diagrs., photos. (NACA RM E51BO2) (Declassified
from Confidential 9/10/54)
An experimental investigation was conducted on the
combustion properties of aluminum as a fuel for use
in jet-powered aircraft. Two techniques of injection
were investigated. In one method aluminum powder
was utilized whereas in the other aluminum wire was
Utilized. Aluminum powder was burned stably and at
combustion efficiencies that averaged about 50 per-
cent, although the thrust obtained was less than that
obtained by burning propane at equivalent condition's*
Aluminum wire was burned with about 75 percent
combustion efficiency. Solid deposits in the comn-
bustors were found to be a serious obstacle to the
use of aluminum as a ram-jet fuel.





NACA RM E51BO8

CORRELATION OF ANALOG SOLUTIONS WITH EX-
PERIIMENTAL SEA-LEVEL TRANSIENT DATA FOR
CONTROLLED TURBINE-PROPELLER ENGINE,
INCLUDING ANALOG RESULTS AT ALTITUDES-
James Lazar and Wilfred L. DeRocher, Jr. August
1951. 36p. diagrs. (NACA RM E51B08) (Declas-
sified from Confidential, 9/10/54)

A satisfactory correlation was obtained between ex-
perimental sea-level transient data and solutions
from the analog representation. The analog repre-
sentation is accomplished by transfer functions that
were formed from a frequency-response analysis of
the experimental transient data as obtained from the
controlled engine. This analog representation was
then used to compute system response at altitude*







NACA
RESEARCH ABSTRACTS NO.71



NACA RM ESID26

AN ALYSIS O F EX PER MENTAL, SEA-LEVEL
TRANSIENT DATA AND ANALOG METHOD OF
OBT AINING A LTIT UDE RESPONSE: FOR TURBINE-
PROPELLER ENuGINE WITTH RELAY-TYPE SPEED
CONTROL. George Vasu and George J. Pack.
May 17, 1951. 28p. diagrs., photo. (NACA
RM E51D26) (Declassified from Confidential,
9'00,54)
Correlation has been established between transient
engine and control data obtained experimentally and
data ,blalnnd by stimulating the engine and control
with an analog computer. This correlation was
established at sea-level conditions for a turbine-
propeller engine with a relay-type speed control.
The behavior of the controlled engine at attitudes
of 20,000 and 35,000 feet was determined with an
analog cuompurer using the altitude pressure and
temperature generalization factors to calculate the
new engine constants for these altitudes. Because
the engine response varies considerably at altitude
some type of compensation appears desirable and
four methods of compensation are discussed.





NACA RM E51Fll

ALTITUDE-IGNITION LIMIT OF A TURBOJET
ENGINE USING A CONDENSER-DISCHARGE
INITION SYSTEM. John C MAms rong.(Dcetober

sifted from Confidential, 9/10/54)






spark and 6 sparks per second.





NACA RM E51Fl8

VELOCITY AND TEMPERATURE FIELDS IN CIR-
CULAR JET EXPANDING FROM CHOKED NOZZLE
INuTO QUIESCENT AIR. MUorris D. Rousso and
Fred D. Kochendorfer. July 1951. 34p. diagrs.,
photos. (NACA RM E51F18. Formerly
RM E50EO3a) (Declassified from Confidenital.
9,'10; 54)
The Mach number and temperature profiles in jets
expanding from convergent and convergent-divergent
nozzles are presented for several values of nozzle-
exit pressure ratio. The effects of jet temperature,
Reynolds number, and humidity on jet spreading are
brlefly evaluated. The results indicated that the
downstream Mach number profiles for a heated jet
a re slightly narrower than those for an unhated jet,
whereas the downstream temperature prof eles were
unaffected by nozzle temperatre change, and that
the effects of Reynolds numer and humidity were
negligible.


NACA RM L51E01

CONTRIBUTIONS OF WIG, TAIL, AND FUSELAGE
TO THE AERODYNAMIC CHARACTERISTICS OF A
SEMISPAN MODEL OF A SUPERSONIC AIRPLANE
CONFIGURATION AT TRANSONIC SPEEDS FROM
TESTS BY THE NACA WING-FLOW METHOD.
Norman S. Silsby and James M. Mnclay. July 1951.
34p. diagrs., photos., tab. (NACA RM L51E01)
(Declassified from. Confidential, 9/10/54)

An investigation has been made by the NACA wing-
flow method at transonic speeds to determine the
contributions of wing, tail, and fuselage to the aero-
dynamic characteristics of a semnispan airplane
model having a long slender fuselage and a straight
wing and tail of low aspect ratio with faired sym-
metrical double-wedge airfoil sections 4l.6-percent-
chord in thickness. Lift, drag, and pitching moments
were obtained for the complete model, wing-fuselage
configuration, fuselage-rall configuration, and
fuselage alone. The M~ach number range of the tests
was from 0.60 to 1.13, and the Reynolds number
range was from about 0.3 x 106 to 0.7 x 106.





NACA RM L51E07

SYSTEM ANALYSES AND AUTOPILOT DESIGN FOR
AUTOMATIC ROLL STABILIZATION OF A SUPER-
SONIC PILOTLESS AIRCRAFT". Jacob Zarovsky.
Juy191. sifi to aC idtab. (NACA RM L51E07)

Automatic roll stabilization system analysis and roll



eernc-druirolautopilot wialdsg av encnuthd ao pass
elpersoicapl ntwork andsitablet gat in djustment ap-

peier (to beaproi acdtic l ad ralpizble mans ofelc


providing the required system roll stabilization
characteristics.




NACA RM L51E09

WING-FLOW STUDY OF PRESSURE-DRAG REDUC-
TION AT TRANSONIC SPEED BY PROJECTIN A
JET OF AIR FROM THE NOSE OF A PROLATE
SPHEROID OF FINENESS RATI 6. Mitchell
Lopatoff. October 1951. 20p. photos., diagrs., tab.
(NACA RM L51E09) (Declassified f rorri Conf adentlal,
9/10/54)

Contains studies by NACA wing-flow method of
pressure-drag reduction obtaied at transonic speed
by projecting a jet of air from the nose of a prolate
spheroid, Comparisons of pressure distributions
with and without the jet are made. ShadowRgraphs of
the model obtained in a small supersonic tunnel at a
constant M~ach numer of 1, 5 while the thrust of the
jet was varied are presented.







NACA
RESEARCH ABSTRACTS NO. 71


Results are presented of an investigation made in the
Langley high-speed 7- by 10-foot tunnel to determine
the high-speed static lateral and directional char-
acteristics of a 1/10-scale model of the X-1 tran-
sonic research airplane from a Mach numer of 0.40
to 0.88.



NACA RM L51FO6a

AEROD)YNAMIIC CHARACTERISTICS AT TRAN-
SONIC SPEEDS OF A TAPERED 450 SWEPTBACK
WING OF ASPECT RATIO 3 HAVING A FULL-SPAN
FLAP-TYPE CONTROL. TRANSONIC-BUMP
METHOD. Vernard E. Lockwood and Joseph E.
Fikes. August 1951. 35p. diagrs. (NACA
RM L51FO6a) (Declassified from Confidential,
9/10/54)


'lph neoe n omoef c ets wre omte ,by te
transonic-bump method on a wing having a quarter
chord sweepback of 45. 58o, aspect ratio of 3, a taper
hai of05dand os NACA 6A01 eto mln a

flections from -270 to 50, and Mach numbers from
06 In o'7.Thheeinvestio htion was made seithethe Tga
results from the bumps, an investigation of a large
model at higher Reynolds number, and available
estimation methods are compared.




NACA RM L51F08

A METHOD FOR THE DESIGN OF SWEPTBACK
WINGS WARPED TO PRODUCE SPECIFIED FLIGHT
CHARACTERISTICS AT SUPERSONIC SPEEDS.
Warren A. Tucker. September1951. 52p. diagrs.,
2 tabs. (NACA RM LS1F08) (Declassified from
Confidential, 9/10/54)

A method is presented for designing sweptback wings
to be self-trimming at a given set of flight conditions.
This characteristic is achieved by warping the wing
in a particular manner, that is, by giving the wing a
certain combination of angle of attack, twist, and
camber. The method applied directly to a wide class
of sweptback wings. The application to any specific
wing is simplified to a routine computational pro-
cedure, and a discussion is given of some points to
be considered in the application to a practical case.
Several illustrative examples are worked out, and
te reuting wings are shown to be feasiable to




NACA RM L51F08a

SMALL-SCALE INVESTIGATION AT TRANSONIC
SPEEDS OF THE EFFECTS OF THICKENING THE
INBOARD SECTION OF A 450 SWEPTBACK WING
OF ASPECT RATIO 4, TAPER RATIO 0.3, AND
NACA 65A006 AIRFOIL SECTION. Kenneth P.
Spreemann and William J. Alford, Jr. August 1951.
21p. diagrs., photo. (NACA RM L51F08a) (Declas-
sified from Confidential, 9/10/54)


NACA RM L51E18

RESULTS OF FLIGHT TESTS TO DETERMINE
DRAG OF PARABOLIC AND CONE-CYLINDER
BODIES OF VERY LARGE FINENESS RATIOS AT
SUPERSONIC SPEEDS. Clement J. Welsh and
Carlos A. deMoraes. August 1951. 17lp. diagrs.,
photos. (NACA RM L51E18) (Declassified from
Confidential, 9/10/54)

Results of free-flight investigation at supersonic
speeds to determine zero-lift drag of bodies of revo-
lution with varying fineness ratios are presented for
both parabolic and cone-cylinder bodies'



NACA RM L51E25a

A PRELIMINARY INVESTIGATION OF COMBUSTION
WITH ROTATING FLOW IN AN ANNULAR COM-
B9U5S1 OCHdAgBsER ph a R. Swartz S ptember
(Declassified from Confidential, 9/10/54)

Aapre Umndat oinvesti kjsnof Hlaon-stablt and

m extur dw o ann ga cmusto chamber wasT cn-

e"""':gmixurburned a h gher a il-nle st n m
obtained with straight-flow burning. Unsteady burn-
ing, accompanied by the pulsations and intense noise
usually present in straight-flow burning, was not
present in the rotating-flow burning. The external
axial-flame length was appreciably less and the
flame divergence was greater with rotating flow.



NACA RM L51F01

COMPARISON OF AIRFOIL SECTIONS ON TWO
TRIANGULAR-WING-FUSELAGE CONFIGURATIONS
AT TRANSONIC SPEEDS FROM TESTS BY THE
NACA WING-FLOW METHOD. Albert W. Hall and
James M. McKay. August 1951. 23p. diagrs.,
photo., tab. (NACA RM L51F01) (Declassified from
Confidential, 9/10/54)

Tests were made by the NACA wing-flow method at
transonic speeds on four triangular-wing-fuselage
models. Two models had wings of aspect ratio 2.31
with NACA 65009 and 9-percent-thick biconvex see-
tions and two models had wings of aspect ratio 4.00
with NACA 65006 and 6-percent-thick double-wedge
sections. Lift, drag, pitching-moment, and angle-
of-attack measurements are presented for a Mach
number range of 0.75 to 1.075. The test Reynolds
number was approximately 1.5 x 106.



NACA RM L51F01a

STATIC LATERAL STABILITY CHARACTERISTICS
OF A 1/10-SCALE MODEL OF THE X-1 AIRPLANE
AT HIGH SUBSONIC MACH NUMBERS. Richard E**
Kuhn and James W. Wiggins. August 1951. 25p.
diagrs., photos. (NACA RM L51F01a) (Declassified
from Confidential, 9/10/54)







NACA
RESEARCH ABSTRACTS NO.71


An investigation was conducted in the Langley high-
speed 7- by 10-fool tunnel over a Mach nuber
range of 0.60 to 1.08 to determine the aerodynamic
effects of thlekenlng mne Inboard 40 percent of a
semispan wing with meS quarter-chord line swept
back 450, an aspect ratio of 4, a taper ratio of 0.3,
and an NACA 64A006 aliroil section. Lift, drag,
pltching moment, and bending moment were obtained
for the twro wings. Also Included in the paper are
some comparisons of experimental results with
theory, corrected to elastic conditions.



NACA RM L51Fl2

TABULATED PRESSURE COEFFICIENTS AND
AERODYNAMIlC CHARACTERISTICS MEASURED IN
F LIGHT ON THE WING OF THE D-558-I RESEARCH
AIRPLANE THROUGH A MACH NUMBER RANGE
OF 0.80 TO 0.89 AND THROUGHOUT THE
NORMAL- FORCE -COE EFFICIENT RANGE AT MLACH
NUMBERS OF 0.61, 0.70, 0.855, AND 0.88. Earl
R. Keener and Rozalia M. Bandish. August 1951.
43p. diagrs.. photos.. 7 tabs. (NACA RM L51F12)
(Declassified froml Confidential, 9/10/54)

Presents tabulated pressure coefficients and aero-
dynamie characteristics obtained in flight from
pressure distrlbutlons over six chordwise rows of
ortfices on a wing of the Douglas D-558-I research
airplane (BuAerro No. 37972). It includes data ob-
tained th roughout a Mlach aumer range of 0.80 to
0.89 and t throughout the no rmal-force-coefficient
range at Mi = 0.61, 0.70, 0.855, and 0.88.



NACA RM L51Fl9

TH EOR ET ICA L ANA LYSES TO DETERMINE UN-
BA LANC ED T RAI LING-EDGE CONTROLS HAVING
M INIM UM HINGE MOMENTS DUE TO DEFLECTION
AT SUPERSONIC SPEEDS. Kennith L. Goin.
November 1952. 52p. dlagrs., tab. (NACA
RM L51F19) (Declasslisedfrom Contidential, 9/10/54)

Theoretical analyses have been made to determine
the plan forms of unbalanced trailing-edge flap-type
cont rols having minimum hinge moments due to de-
flectlon and requiring minimum work to overcome
the hange mloments due to deflection at supersonic
speeds. Ratios of h~it and rolling moment to hinge
moment and hItI and rolling moment to deflection
work at fIxed values of lift and rolling effectiveness
were used as bases for the analyses. The effects of
control plan form, control location, and Mach num-
ber have been considered.



NACA RM L51F26

EF FECTS OF PROPELLER-SHANK GEOMETRY
AND PROPELLER-SPINN ER-JUNCTURE CONFIG-
SU R ION OOW GA R CdERITC C F ANb N A 1
AN EIGHT-D3LADE DUA L-ROTATIONr PROPELLER.
Arvld L. Kielth, Jr., Gene J. Bingham and Arnold J.
Rab n. ( ptemlber l51 F 6) edliasgrs, photos., 5
Conisdentlal, 9/10/~54)


Results of low-speed investigation of aerodynamic
characteristic of NACA 1-series cowling-spinner
combination equipped with several eight-blade dual-
rotation propellers differing in shank-thickness
ratio are presented. Studies of sealed and faired
propeller spinner juncture~s, four junctures that
permit blade rotation, and two methods for retarding
spinner boundary-layer separation are included.



NACA RM L51F28

EFFECTS OF SPANWISE THICKNESS VARIATION
ON THE: TRANSONIC AERODNAMIC CHARACTER-
ISTICS OF WINGS HAVING 350 OF SWEEPBACK,
ASPECT RATIO 4, AND TAPER RATIO 0.60.
William D. Morrison, Jr. and Paul G. Fournier,
September 1951. 21p. diagrs., photo. (NACA
RM L51F28) (Declassified from Confidential,
9/10/54)

The aerodynamic characteristics of a wing tapered
in thickness from 6 percent at the root to 2 percent
at the tip having 350 sweepback, aspect ratio 4, and
taper ratio 0.60 have been determined by a reflection
plane technique over a Mach numer range from 0.60
to 1.08 at a Reynolds number of about 650,000. The
results of this investigation are compared with those
of a wing of identical plan form but of a constant
thickness ratio. Theoretical subsonic and low-
supersonic calculation of lift-curve slope, aero-
dynamic center, and lateral center-of-lift location
are compared with experimental results.




NACA RM L51F29

BASE PRESSURES MEASURED ON SEVERAL
PARABOLIC-ARC BODIES OF REVOLUTIO IN
FREE FLIGHT AT MACH NUMBERS FROM 0.8 TO
1.4 AND AT LARE: REYNOLDS NUMBERS. Ellis
Katz and William E. Stoney, Jr. October 1951. 20p.
diagrs., photos. (NACA RM L51F29) (Declassified
from Confidential, 9/10/54).

Base pressures were measured on several fin stabi-
lized bodies of parabolic-are profile in free flighrt at
Mach numbers from 0.8 to 1.4 and at Reynolds num-
bers from 20 to 130 million. The bodies varied in
length from 6 to 25 diameters and had afterbodies
which converged to base areas equal to 19.1 percent
of the frontal areas. Pressures were also measured
on the side of the bodies immediately ahead of the
bases.




NACA RM L51GO3

AN INVESTIGATION AT TRANSONIC SPEEDS OF
THIE EFFECTS OF CONTROL CHORD AND SPAN

O PTEHRE OIME P H A GR F ASPS C RATIO
2.5. TRANSONIC-BUMP METHOD. Raymond D.
Vogler, Vernard E. Lockwood and Thomas R.
Tune5. G ) ambrs f91 d6p f an n(NACA
9/10 54)







NACA
RESEARCH ABSTRACTS NO.71




NACA RM LS1HO2

A HEREIA INETGTO F QT HE NLU -
UPON THE DYNAMIC PERFORMANCE CHARAC-
TERISTICS OF A SUPERSONIC CANARD MIISSILE
CONFIGURATION WITH A P~ITCH-ATTITUDE CON-
TROL SYSTEM. Anthony L. Passera. October
1951. 32p. diagrs., photos., 2 tabs. (NACA
RM L51HO2) (Declassified from Confidential,
9/10/54)

A theoretical investigation is made to determine the
effects of the natural frequency of a single-degree-
of-freedom autopilot upon the dynamic performance
characteristics of a supersonic canard missile con-
figuration with rate-damping and pitch-attitude con-
trol system. These effects are presented in the
form of pitch-angle, control-surface deflection, and
normal acceleration transients for several flight
emanditi nsain response to a unit step attitude com-





NACA RM L51H16a

LOW-SPEED WIND-TUNNEL INVESTIGATION OF A
FIXED AND A FREE-FLOATING WING-TIP AILER-
ON ON A WING WITH LEADING EDGE: SWEPT
BACK 51.30. R.G. MacLeod. February 1952.
17p. diagrs., photo., tab. (NACA RM L51H16a)
(Declassified from Confidential, 9/10/54)

Contains results of a low-speed wind-tunnel investi-
gation of a fixed and a free-floating wing-tip aileron
on a wing with leading edge swept back 51.30. Re-
sults indicate that the rolling effectiveness of both
ailerons was maintained up to very large angles of
attack while the yaw due to aileron deflection was
more favorable for the free-floating aileron.






NACA RM L51H22

AERODYNAMIC CHARACTERISTICS AT TRANSONIC
SPEEDS OF A 600 DELTA WING EQUIPPED WITH A
CONSTANT-CHORD) FLAP-TYPE CONTROL WITH
AND WITHOUT AN UNSHIELDED HORN BALANCE.
TRANSONIC-BUMP METHOD. Harleth G. Wiley
and Leon Zontek. September 1952. 25p. diagrs.
(NACA RM L51H22) (Declassified from ConfidentiaL.
9/10/54)

This paper presents the aerodynamic characteristics
of a 600 delta wing of aspect ratio 2.31, taper ratio
0, and an NACA 65-006 airfoil section, which wras
equipped with a constant-chord control, with and
without an unshielded triangular horn balance. Lift,
pitching moment, and hinge moment were obtained at
various angles of control deflection and angles of
attack for both types of controls through a Mach num-
ber range of 0.6 to 1.18.


This paper presents the results of a control inves-

o ario :.s spnw demg ni ) es t5 on r s
ratio of 0.625. Rolling and pitching moments and
lift of the semispan wing-fuselage combination were
obtained through a Mach range of 0.6 to 1.18 and at a
Reynolds number near 1.2 x 106. The control deflec-
tions ranged from Go to 150. The investigation was
conducted in the Langley high-speed 7- by 10-foot
tunnel using the transonic bump.







NACA RM L51GO9

PRELIMINARY INVESTIGATION OF THE EFFECTS
OF RECTANGULAR VORTEX GENERATORS ON
THE PERFORMANCE OF A SHORT 1.9:1
STRAIGHT-WALL ANNULAR DIFFUSER. Charles
C. Wood. October 1951. 27p. diagrs., photo., tab.
(NACA RM L51GO9) (Declassified from Confidential,
9/10/54)

Results are presented of a preliminary investigation
of vortex generators introduced into the region of
boundary layer near the inlet of a short annular dit-
fuser to increase turbulent mixing in an attempt to
prevent separation and resulting adverse effects.
Rectangular noncambered airfoils were used as vor-
tex generators and were varied in chord, span, angle
of attack, number, and location. Some of the vortex
generator arrangements resulted only in small in-
creases to diffuser performance, while other ar-
rangements resulted in substantial increases. The
effect of one of the better vortex-generator arrange-
ments was to eliminate separation of the flow en-
tirely and to increase the diffuser effectiveness 17
percent.








NACA RM L51G20

A CORRELATION OF EXPERIMENTAL ZERO-LIFT
DRAG OF RECTANGULAR WINGS WITH SYMMET-
RICAL NACA 65-SERIES AIRFOIL SECTIONS BY
MEANS OF THE TRANSONIC SIMILARITY LAW
FOR WINGS OF FINITE ASPECT RATIO. Edward
C. B. Danforth. September 1951. 20p. diagrs.
(NACA RM L51G20} (Declassified from Confidential,
9/10/54)

This paper contains a correlation of zero-lift tran-
sonic pressure drag of rectangular wings of finite
aspect ratio by means of a transonic similarity law
derived in NACA TN 2273. The data correlated were
for wings of exposed aspect ratio between 1.12 and
6.25 and thickness ratios between 0.03 and 0.12, all
with symmetrical NACA 65-series airfoil sections.





NACA
RESEARCH AEISTRACTS NO. 71

NACA RM L51H23

THE E EFFECTS ON THE AERODYNAMIC CHARAC-
TERISTICS OF REVERSING THE WING OF A TRI-
ANGULAR WING-BODY COMBINATION AT TRAN-
SONIC SPEEDS AS DETERMINED BY THE NACA
WING-FLOW METHOD. James M. McKay and
Albert W. Hall. October 1951. 22p. diagrs., photo.>
2 tabs. (NACA RM L51H23) (Declassified from
Confidential, 9/10/54)

Tests were made by the NACA wing-flowr method on
two triangular wing-fuselage models of low aspect
ratio with 6i-percent-thick biconvex sections. On one
model the apex angle of the wing was forward and on
the other the apex angle was trailing. Lift, drag,
pitching moment, and angle-of-attack measurements
are presented for a Mach number range of 0.75 to
1.075. The test Reynolds number was approximately
1.5 x 1016*





NACA RM L51H27

INVESTIGAION OF WING-TIP AILERONS ON A
51.30 SWEPTBACK WING AT TRANSONIC SPEEDS
BY THE TRANSONIC-BUMP METHOD. William C.
Moseley, Jr. and James M. Watson. Nlovemnber 1951.
60p. dlagrs. (NACA RM L51H27) (Declassified
from Confidential, 9/10/54)

Three wing-tip ailerons were tested through the .
Iransomec range on a 51.30 sweptback wing. One was
a triangular wing tip deflectable about an axis normal
to the leading edge of the wing; the second was a tri-
angular tip added to the basic wing; and the third,
utihzlng the extended tip area, was obtained by de-
flecting the area aft of a spanwise axis through the
0.50-tap-hord station of the basic wing. The last
tw~o allerons provided lateral control over the Mach
number range at lorw angles of attack.



NACA RM L51H28

AERODY NAMIC CHARACTERISTICS AT TRANSONIC
SPEEDS OF A WIN HAVING 450 SWEEP, ASPECT
RATIO 8, TAPER RATIO 0.45, AND AIRFOIL
SECTIONS VARYING FROM THE NACA 63A010
SECTION AT THE ROOT TO THE NACA 63A006
SECTION AT THE TIP. Wrilliam D. Morrison, Jr.
and Paul G. Fournier. January 1952. 22p. diagrs.,
pholo. (NACA RM L51H28) (Declassified from
Confidential, 9/10/54)

The aerodynamic characteristics of a wing tapered
In thickness ratio from 10 percent at the root chord
to 6 percent at the tip chord, having 450 sweepback,
aspect ratio 8, and taper ratio 0.45 have been deter-
maned by a reflection-plane technique over a Mach
number range? from 0.60 to 1.05 at a Reynolds number
of about 500,000. The results of this investigation
are compared with those obtained for a wing of iden-
tlcal plan form but of a constant 12-percent section
thlcckness. Theoretical subsonic calculations of lift-
curve slope, and aerodynamic-center and lateral-
center-of-lift locations are compared with the exper-
imental results.


NACA RM L51H29

THE EFFECT OF RAKING THE AILERON TIPS ON
THE LATERAL-CONTROL AND HINGE-MOMIENT
CHARACTERISTICS OF A 20-PERCENT-CHORD
PARTIAL-SPAN OUTBOARD AILERON ON A WING
WITH LEADING EDGE SWEPT BACK 51.30.
Alexander D. Hammond. November 1951. 41p.
diagrs., photo., tab. (NACA RM L51H29) (Declas-
sified from Confidential, 9/10/54)

A wind-tunel investigation at low speed was made to
determine the lateral-control and hinge-ntoment
characteristics of an unsealed plain-radius-nose 20-
percent-chord flat-sided partial-span outboard ailer-
on having various plan forms on a wing with the lead-
ing edge swept back 51.30, aspect ratio 3.06, taper
ratio 0.49, and NACA 651-012 airfoil sections per-
pendicular to the 55.6-percent-chord line, The re-
aults of the investigation indicate that, for ailerons
having the same area, plan form has little or no
effect on the variation of rolling-moment coefficient
with aileron deflection. Changes in the aileron plan
form, however, have an appreciable effect on the
hinge-moment parameter Ch6'




NACA RM L51H30

SUMMARY OF RESULTS OBTAINED BY
TRANSONIC-BUMIP METHOD ON EFFECTS OF
PLAN FORM AND THICKNESS OF LIFT AND DRAG
CHARACTERISTICS OF WINGS AT TRANSONIC
SPEEDS. Edward C. Polhamus. November 1951.
33p. diagrs., tab. (NACA RM LS1H30) (Declas-
sified from Confidential, 9/10/54)

This paper presents a summary of the effects of plan
form and thickness on the lift and drag characteris-
tics of wings at transonic speeds and comparisons
with subsonic, transonic, and supersonic theories.
The data considered in this summary were obtained
during a transonic research program conducted in
the Langley high-speed 7- by 10-foot tunnel by the
transonic bump method. The Reynolds numbers
of the tests were generally less than 1 x 106,


NACA RM L51I06

EFFECTS OF HORIZONTAL-TAIL POSITION, AREA,
AND ASPECT RATIO ON LOW-SPEED STATIC
LONGITUDINAL STABILITY AND CONTROL CHAR-
ACTERISTICS OF A 600 TRIAINGULAR-WIN
MODEL HAVIN VARIOUS TRIANGULAR-ALL-
MOVABLE HORIZONTAL TAILS. Byron M. Jaquet.
December 1951. 61p. diagrs., photo., tab. (NACA
RM L51I06) (Declassified from Confidential,
9/10/54)

A low-speed investigation (Mach number of 0.17 and
Reynolds number of 2.06 x 106) was made in the
Langley stability tunnel to determine the Static longi-
tudinal stability and control characteristics of a 600
triangular-wing mode having various all-movable
triangular horizontal tails. Effects of tail length,
height, area, and aspect ratio are presented. All-
movable tails are compared with constant-chord
flaps and half-delta tip controls.








NACA
RESEARCH ABSTRACTS NO. 71




speed 7- by 10-foot tunnel over a Mach number range
of 0.59 to 1.10 to determine the effects of twist and
chamber on the aerodynamic characteristics of a
sweptback wing rotated from its original design posi-
tion. This paper presents the results of the investi-
gation of the two semispan wings swept back 600 42';
one wing was untwisted and uncambered, the other
incorporated twist and camber. The wings had
aspect ratios of 1. 94, taper ratios of 0.44, and
NACA 64A-series airfoil sections tapered in thick-
ness. Lift, drag, pitching moment, and handing
moment were obtained for the two wings.


NACA RM L51I21

SMALL-SCALE INVESTIGATION OF THE EFFECTS
OF TWIST AND CAMBER ON THE AERODYNAMIC
CHARACTERISTICS OF A 600 42' SWEPTBACK
WING OF ASPECT RATIO 1.94. Kenneth P.
Spreemann and William J. Alford, Jr. January 1952.
19p. diagrs., photo., tab. (NACA RM L51I21)
(Declassified from Confidential, 9/10/54)

An investigation of two semispan wings having the
same plan form was conducted in the Langley high-


NACA-Langley 10-7-54 4M






































































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