Research abstracts and reclassification notice

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Title:
Research abstracts and reclassification notice
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NACA research abstracts and reclassification notice
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37 v. : ; 27 cm.
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English
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United States -- National Advisory Committee for Aeronautics
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National Advisory Committee for Aeronautics
Place of Publication:
Washington, D.C
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completely irregular

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Aeronautics -- Abstracts -- Periodicals   ( lcsh )
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abstract or summary   ( marcgt )
federal government publication   ( marcgt )

Notes

Dates or Sequential Designation:
No. 94 (Jan. 11, 1956)-no. 130 (Sept. 30, 1958).
General Note:
Title from caption.
Statement of Responsibility:
National Advisory Committee for Aeronautics.

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University of Florida
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Resource Identifier:
aleph - 001469325
notis - AGY1018
oclc - 14184154
lccn - 86657026
sobekcm - AA00005288_00002
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AA00005288:00002

Related Items

Preceded by:
Research abstracts
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Publications announcements

Full Text
y3.l/ 21 ; IL4

National Advisory Committee For Aeronautics

Research Abstracts

and
Reclassification Notice DEC 10 i957

NO. 122 DECEtB 3, 1

SPECIAL NOTICE


All NACA publications that were issued under the following series designatdSons-'a nbw-
declassified:
NACA ACR (Advance Confidential Report)
NACA ARR (Advance Restricted Report)
di ir .,I l, ll i


NALCA
NACA
NACA


LCD
RB
MR


SI.U Ll[ llll d.l DUll1 LIII J
(Restricted Bulletin)
(Memorandum Report)


.j

. "I I 'k I


None of these series designations is currently used; the last publications designated by them were
issued in 1947.

All NACA Research Memorandums with 1946 code numbers (such as E6B12) are now declassified.


The declassification of each publication
present announcement is made for the
individual announcements.


CURRENT
NACA RESEARCH REPORTS

NACA Rept. 1309
AERODYNAMIC CHARACTERISTICS AT HIGH
SPEEDS OF RELATED FULL-SCALE PROPELLERS
HAVING DIFFERENT BLADE-SECTION CAMBERS.
Julian D. Maynard and Leland B. Salters, Jr. 1957.
ii, 24p. diagrs., photos., tab. (NACA Rept. 1309.
Supersedes RM L8E06)
Comparisons are made of results obtained In wind-
tunnel tests of related full-scale propellers over a
range of blade angles from 200 to 550 at airspeeds
up to 500 miles per hour to evaluate the combined
effects of blade-section camber and compressibility
on propeller aerodynamic characteristics.


NACA RM E57G11
ACCELERATION IN FIGHTER-AIRPLANE
CRASHES. Loren W. Acker, Dugald O. Black, and
Jacob C. Moser. November 1957. 78p. diagrs.,
photos. (NACA RM E57Gll)
Full-scale crashes were conducted with FH-1 jet
fighter airplanes under circumstances approximating
those observed in the military service. These
crashes simulated unflared landings at impact angles
of 180, 220, and 270, a ground cart wheel, and a
S ground loop. The magnitude, duration, and direc-
-b tion of the crash accelerations were measured on the
U airplane structure and on an anthropomorphic dummy
installed in the cockpit. The accelerations meas-
ured are compared with existing data on human tol-
erance to the sudden loads that occur in crashes to
see whether the human tolerance had been exceeded.


in these groups has been previously announced. The
benefit of those who may not have access to all of the


NACA RM E57G16a
THEORETICAL ROCKET PERFORMANCE OF
JP-4 FUEL WITH SEVERAL FLUORINE-OXYGEN
MIXTURES ASSUMING FROZEN COMPOSITION.
Sanford Gordon and Kenneth S. Drellishak.
November 1957. 62p. diagrs., tabs.
(NACA RM E57G16a)
Data were calculated for a chamber pressure of
600 pounds per square inch absolute, fluorine in
oxidant from 0 to 100 percent, pressure ratios of
1 to 1500, and a range of equivalence ratios.
Parameters included are specific unpulse, com-
bustion and exit temperatures, molecular weight,
characteristic velocity, coefficient of thrust, area
ratio, specific heat, isentropic exponent, viscosity,
thermal conductivity, and combustion composition.





NACA TN 4046
A COMPARATIVE ANALYSIS OF THE PERFORM-
ANCE OF LONG-RANGE HYPERVELOCITY
VEHICLES. Alfred J. Eggers, Jr., H. Julian Allen,
and Stanford E. Neice. October 1957. (i), 66p.
diagrs. (NACA TN 4046. Supersedes RM A54L10)
A simplified analysis is made of the motion and aero-
dynamic heating of long-range ballistic-, skip-, and
glide-type vehicles. The ballistic vehicle appears
relatively attractive because convective heat transfer
can be reduced by using blunt shapes. The glide
vehicle appears attractive because it has a relatively
efficient trajectory, and the possibility of substan-
tial radiative cooling. These vehicles compare fav-
orably, in the sense of the Brequet range equation,
to the supersonic airplane for very long-range flight.


-AVAILABLE ON LOAN ONLY


ADDRESS REQUESTS FOR DOCUMENTS TO NACA. 1512 HST
ABOVE EACH TITLE. AND THE AUTHOR


. NW, WASHINGTON 25 D. C. ITING CODE NUMBER WHICH IS










CURRENT NACA RESEARCH REPORTS


NACA TN 4047

A STUDY OF THE MOTION AND AERODYNAMIC
HEATING OF MISSILES ENTERING THE EARTH'S
ATMOSPHERE AT HIGH SUPERSONIC SPEEDS.
H. Julian Alien and A. J. Eggers, Jr. October
1957. 61p. diagrs. (NACA TN 4047. Supersedes
RM A53D28)

A simplified analysis of the velocity and deceleration
history of missiles entering the earth's atmosphere
at high supersonic speeds is presented. The results
of this motion analysis are employed to indicate
means available to the designer for minimizing aero-
dynamic heating. The heating problem considered
involves not only the total heat transferred to a mis-
sile by convection, but also the maximum average
and local time rates of convective heat transfer.


NACA TN 4048

MOTION OF A BALLISTIC MISSILE ANGULARLY
MISALINED WITH THE FLIGHT PATH UPON EN-
TERING THE ATMOSPHERE AND ITS EFFECT UP-
ON AERODYNAMIC HEATING, AERODYNAMIC
LOADS, AND MISS DISTANCE. H. Julian Allen.
October 1957. 66p. diagrs., tabs. (NACA TN 4048.
Supersedes RM A56F15)

An analysis is given of the oscillating motion of a
ballistic missile which upon entering the atmosphere
is angularly misalined with respect to the flight path.
The history of the motion for some example missiles
is discussed from the point of the effect of the motion
on the aerodynamic heating and loading. The miss
distance at the target due to misalinement and to
small accidental trim angles is treated. The stabil-
ity problem is also discussed for the case where the
missile is tumbling prior to atmospheric entry.


NACA TN 4087

DROP-SIZE DISTRIBUTION FOR CROSSCURRENT
BREAKUP OF LIQUID JETS IN AIRSTREAMS.
Robert D. Ingebo and Hampton H. Foster.
October 1957. 36p. diagrs., photos, tabs.
(NACA TN 4087)

Photographic and sampling techniques were combined
to obtain drop-size data for ranges of injector,
liquid, and airstream variables. The following
empirical expression correlated the ratio of volume-
median diameter D30 to orifice diameter Do with
Weber-Reynolds number ratio
D30/Do = 3.9(We/Re)0-25 where We = o/psDoVs,
Re = DoVs/v, a and v are surface tension and
kinematic viscosity, respectively, of the liquid; and
Vs and ps are free-stream velocity and density,
respectively, of the air. A drop-size-distribution
equation based on maximum observed drop diameter
and Weber-Reynolds number ratio was also derived.

dR- W06( 024 D5 -22.3(We/Re)0-04D/Dmax
dD 10\Re) 6 e
Dmax


NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122


NACA TN 4097

INVESTIGATION OF SOME MECHANICAL PROP-
ERTIES OF THERMENOL COMPRESSOR BLADES.
Donald F. Johnson. October 1957. 14p. diagrs.,
photo. (NACA TN 4097)

A series of tests were made comparing the mechani-
cal properties of similar compressor blades of AISI
type 403 stainless steel and thermenol. Eighth-stage
J47 and J65 compressor blades of each material were
tested. Modulus of elasticity, modulus of rigidity,
and damping were slightly lower for thermenol.
However, thermenol showed better resistance to
corrosion by sea water, and the thermenol blades
proved equal or superior to the stainless-steel blades
in fatigue strength.







NACA TN 4124

EFFECT OF GROUND PROXIMITY ON THE AERO-
DYNAMIC CHARACTERISTICS OF A FOUR-ENGINE
VERTICAL-TAKE-OFF-AND-LAN DING
TRANSPORT-AIRPLANE MODEL WITH TILTING
WING AND PROPELLERS. William A. Newsom,
Jr. October 1957. 15p. diagrs., photo., tab.
(NACA TN 4124)

An investigation was conducted on an airplane model
with the wing at an angle of incidence of 900 for a
range of heights of the model above the ground and
included force tests and tuft studies of the flow field
caused by the propeller slipstream. The results
indicate that, when the model was hovering near the
ground, there was a strong upwash in the plane of
symmetry which caused both an increase m lift be-
cause of the up load on the fuselage and an increase
in the nose-down pitching moment because the rear
part of the fuselage was longer than the front part.







NACA TN 4129

ANALYSIS OF OPERATIONAL AIRLINE DATA TO
SHOW THE EFFECTS OF AIRBORNE WEATHER
RADAR ON THE GUST LOADS AND OPERATING
PRACTICES OF TWIN-ENGINE SHORT-HAUL
TRANSPORT AIRPLANES. Martin R. Copp and
Walter G. Walker. November 1957. 18p. diagrs.,
tabs. (NACA TN 4129)

Airspeed, altitude, and acceleration data were ob-
tained from transports flown by one airline before
and subsequent to the installation of airborne weather
radar. A comparison of the results indicated that
the magnitudes of the largest gust velocities and gust
accelerations experienced for a given number of
flight miles during radar operations were approxi-
mately 25 percent less than those experienced before
the radar equipment was installed.


*AVAILABLE ON LOAN ONLY






NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122 3


CURRENT NACA RESEARCH REPORTS



NACA TN 4130

NACA 65-SERIES COMPRESSOR ROTOR PERFORM-
ANCE WITH VARYING ANNULUS-AREA RATIO,
SOLIDITY, BLADE ANGLE, AND REYNOLDS NUM-
BER AND COMPARISON WITH CASCADE RESULTS.
Wallace M. Schulze, John R. Erwin, and George C.
Ashby, Jr. October 1957. 62p. diagrs., photos.,
tab. (NACA TN 4130. Supersedes RM L52L17)

An axial-flow compressor rotor was tested at low
speed and without guide vanes or stators in a 28-inch
test compressor. The rotor was designed for a free-
vortex type blade loading using NACA 65-series
10-percent-thick airfoil sections. In these tests,
the effects of changes of blade-setting angle, annulus-
area ratio, solidity, and Reynolds number on rotor
performance were determined for a range of quantity
flows. Rotor data are compared with estimations
based on uncorrected and corrected cascade data.



NACA TN 4132

FATIGUE INVESTIGATION OF FULL-SCALE
TRANSPORT-AIRPLANE WINGS. VARIABLE-
AMPLITUDE TESTS WITH A GUST-LOADS
SPECTRUM. Richard E. Whaley. November 1957.
43p. diagrs., photos., tabs. (NACA TN 4132)

Crack-initiation areas, frequency of occurrence of
cracks, crack propagation, lifetime to crack initia-
tion of all cracks, lifetime to final failure, and
spread in lifetime are investigated and compared
with results from constant-amplitude tests on simi-
lar wings. Information on X-ray techniques to de-
termine the presence of cracks in hidden elements
is also included.



NACA TN 4143

DEVELOPMENT OF A PISTON-COMPRESSOR TYPE
LIGHT-GAS GUN FOR THE LAUNCHING OF FREE-
FLIGHT MODELS AT HIGH VELOCITY. A. C.
Charters, B. Pat Denardo, and Vernon J. Rossow.
November 1957. (1), 95p. diagrs., photos., tabs.
(NACA TN 4143. Supersedes RM A55GI lI

A light-gas gun has been developed at the Ames
Aeronautical Laboratory to launch small models for
aerodynamic tests. The design of the gun and the
analysis of its performance are presented. The re-
sults of the initial firing trials are discussed. The
firing trials showed good agreement between meas-
ured and predicted velocities and pressures and a
velocity of 15.400 ft/sec was reached. It is con-
cluded that the gun is a satisfactory launcher for
high-velocity free-light tests.


NACA TN 4156

EFFECT OF INITIAL MIXTURE-TEMPERATURE ON
BURNING VELOCITY OF HYDROGEN-AIR MIX-
TURES WITH PREHEATING AND SIMULATED
PREBURNING. Sheldon Heimel. October 1957.
23p. diagrs., tabs. (NACA TN 4156)

Laminar burning velocities were determined in the
temperature range 3000 to 7000 K from schlieren
photographs of open flames. The temperature was
raised in two ways: (1) by preheating of the
hydrogen-air mixtures and (2) by simulated adiabatic
preburning of part of the hydrogen in air at 3000 K so
that initial temperatures of 6000 and 7000 K would be
attained for the resulting mixtures of hydrogen, air,
water vapor, and nitrogen. The temperature de-
pendence of burning velocity was determined for both
preheated and preburned mixtures. With 29.6 and
45.0 percent hydrogen (original mixture), mole-for-
mole substitution of nitrogen for water vapor in the
preburning experiments caused no discernible change
in burning velocity.






NACA TN 4162

STUDY OF SOME BURNER CROSS-SECTION
CHANGES THAT INCREASE SPACE-HEATING
RATES. Donald R. Boldman and Perry L.
Blackshear, Jr. November 1957. 38p. diagrs.,
photos., tab. (NACA TN 4162)

A two-dimensional glass-walled combustor fed with a
homogeneous combustible mixture was used to study
the effect of area blockage on heat release. The
blockage at a single V-gutter flame holder was varied
from 12.5 to 75 percent with negligible effect on com-
bustion efficiency. When a small flame holder was
used and a 62.5-percent restriction introduced down-
stream, the heat-release rate underwent a three- to
fourfold increase. Some effects of downstream
blockage shapes are given.






NACA TN 4167

A RAPID METHOD FOR PREDICTING ATTACHED-
SHOCK SHAPE. Eugene S. Love and Ronald H.
Long. October 1957. 34p. diagrs., tab.
(NACA TN 4167)

A method is presented for the rapid prediction of the
shape of attached shocks emanating from smoothly
contoured axisymmetric and two-dimensional nose
shapes. From a practical viewpoint the accuracy of
the method is comparable to that of the method of
characteristics.


*AVAILABLE ON LOAN ONLY.









CURRENT NACA RESEARCH REPORTS






NACA TN 4169

ATMOSPHERIC TEMPERATURE OBSERVATIONS
TO 100,000 FEET FOR SEVERAL CLIMATOLOGI-
CAL REGIONS OF THE NORTHERN HEMISPHERE.
H. B. Tolefson. November 1957. 26p. diagr., tab.
(NACA TN 4169)

Radiosonde measurements of upper-air temperatures
taken over a 5-year period at nine stations in the
northern hemisphere are summarized in tabular
form in order to provide information on the temper-
atures likely to be encountered during airplane and
missile operations up to 100,000 feet. The results
indicate that the scatter in the temperatures about
the mean generally decreased with increasing alti-
tude from the tropopause to 100,000 feet. Little, if
any, effect of location upon the temperature was ap-
parent for altitudes above about 90,000 feet.






NACA TN 4171

SOME GROUND MEASUREMENTS OF THE FORCES
APPLIED BY PILOTS TO A SIDE-LOCATED AIR-
CRAFT CONTROLLER. Roy F. Brissenden.
November 1957. 17p. diagrs., photos., tab.
(NACA TN 4171)

Ground tests have been made to determine pilots'
force capabilities on a proposed side-located aircraft
controller. In addition to maximum force capabili-
ties, pilots participating in the tests indicated forces
that they considered desirable for operation of the
controller. A neutral position and a range of deflec-
tions and forces are suggested.






NACA TN 4174

WIND-TUNNEL INVESTIGATION OF THE STATIC
LATERAL STABILITY CHARACTERISTICS OF
WING-FUSELAGE COMBINATIONS AT HIGH SUB-
SONIC SPEEDS. TAPER-RATIO SERIES. James
W. Wiggins and Paul G. Fournier. October 1957.
25p. diagrs., photos. (NACA TN 4174. Supersedes
RM L53B25a)

This paper presents the aerodynamic characteristics
in sideslip of wing-fuselage combinations. The
wings have a sweep angle of 450 at the quarter-chord
line, aspect ratio of 4, taper ratios of 0.3, 0.6, and
1.0, and an NACA 65A006 airfoil section. The tests
covered a Mach number range from 0.4 to 0.95, a
Reynolds number range from 1.7 x 106 to 3.2 x 106,
and an angle-of-attack range from -30 to 24o.


NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122

NACA TN 4175

INVESTIGATION OF DEFLECTORS AS GUST ALLE-
VIATORS ON A 0.09-SCALE MODEL OF THE BELL
X-5 AIRPLANE WITH VARIOUS WING SWEEP
ANGLES FROM 200 TO 600 AT MACH NUMBERS
FROM 0.40 TO 0.90. Delwin R. Croom and Jarrett
K. Huffman. November 1957. 28p. diagrs.
(NACA TN 4175)

An investigation has been made in the Langley high-
speed 7- by 10-foot tunnel to determine the effective-
ness of deflectors as gust alleviators on a 0.09-scale
model of the Bell X-5 research airplane with wing
sweep angles from 200 to 600. Results of this study
have indicated that a deflector arrangement can be
used as a gust-alleviation device on swept wings.


BRITISH REPORTS


N-50049*

Royal Aircraft Establishment (Gt. Brit.)
SOME METHODS OF EVALUATING IMPERFECT
GAS EFFECTS IN AERODYNAMIC PROBLEMS.
G. A. Bird. January 1957. 31p. diagrs. RAE
Tech.Note Aero 2488. (Ask for N-50049 )

Simple numerical and graphical procedures are de-
scribed for the calculation of the.imperfect gas ef-
fects on the properties of steady and unsteady one-
dimensional isentropic flows, the Prandtl-Meyer ex-
pansion round a corner and normal and oblique shock
waves. The fundamental equations of each type of
flow have been put into a form in which they may be
solved using the published tables of the equilibrium
properties of gases. Both thermal and caloric im-
perfections have been taken into account but relaxa-
tion time effects have been neglected. Numerical
examples are given for each type of flow although the
main emphasis has been placed on the methods rather
than on the results. These basic methods have been
used to calculate the magnitude of the imperfect gas
effects on a number of specific aerodynamic prob-
lems which have been considered in detail.

N-50128*

Royal Aircraft Establishment (Gt. Brit.)
METHODS FOR DETERMINING THE WAVE DRAG
OF NON-LIFTING WING-BODY COMBINATIONS.
L. M. Sheppard. April 1957. 39p. diagrs. RAE
Aero 2590. (Ask for N-50128*)

The area rule, moment of area rule, and transfer
rule methods for estimating the wave drag of wing-
body combinations are discussed. It is pointed out
that the moment of area rule and the transfer rule are
different forms of the area rule, and that the transfer
rule expresses the interference wave drag in a sim-
ple form. The existing methods of wave drag esti-
mation are restricted to combinations with bodies
having continuous surface slope and here an exten-
sion to combinations with bodies having discontinu-
ous surface slope is given. This paper is concerned
with the theoretical methods and their associated nu-
merical techniques and no numerical results or par-
ticular applications are presented.


*AVAILABLE ON LOAN ONLY.






NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122 5


BRITISH REPORTS



N-50129*

Ministry of Supply (Gt. Brit.)
THE SULF[NUZ TREATMENT OF S.80 AND S.99
STEELS, EFFECT ON FATIGUE STRENGTH AND
OTHER PROPERTIES. July 1957. 15p. diagrs.,
photos., tabs. MOSS & T Memo. 9/57.
(Ask for N-50129*)

Tensile, Izod impact, fatigue and corrosion tests
have been carried out on two steels treated by the
"SULFINUZ" process. A significant improvement
in fatigue strength is noted when compared with
steels treated in a neutral salt bath. Under severe
corrosive conditions the "SULFINUZ" treated speci-
mens show greater attack with marked local action.
A preliminary metallographic examination is
described.


N-50131*

Royal Aircraft Establishment (Gt. Brit.)
THE SHRINKAGE AND EXTENSIBILITY OF HEATED
YARNS. PART I. J. E. Swallow. March 1957.
22p. diagrs. RAE Tech.Note Chem. 1302.
(Ask for N-50131')

The changes in length which occur when nylon,
Terylene and Fortisan yarns are heated, and the
reversibility of these changes, are described. The
effects of stress on yarns which have been heated
are considered. It is shown that stressing may
reduce the dimensional stability conferred by heat-
setting.



N-50132*

Royal Aircraft Establishment (Gt. Brit.)
FACTORS AFFECTING THE THERMAL STABILITY
OF POLYESTER GLASS-FIBRE LAMINATES.
B. A. Blythe and W. W. Wright. April 1957. 20p.
diagrs., tabs. RAE Tech.Note Chem. 1306.
(Ask for N-50132*)

The effects of variation of glass cloth, glass-cloth
treatment, and initiator for the polymerization reac-
tion on the flexural strength at elevated tempera-
tures of polyester glass-fiber laminates have been
studied. The strengths of high- and low-alkali
glass-cloth laminates were comparable. The use of
treated glass cloths resulted in improved initial
strength and strength retention with time; this was
especially true of the Garan finish. Variation of
the initiator used for the curing reaction, although
resulting in a wide variation of the flexural strength
values after particular times at temperature, ap-
peared to have little effect on the characteristics of
the thermal breakdown. The variations in flexural
strength of laminates made using 10 heat-resistant
laminating resins were determined with time at ele-
vated temperature. Very good results were ob-
tained with a phenolic resin.


N-50134*

Royal Aircraft Establishment (Gt. Brit.)
THE GEOMETRY OF WING SURFACES GENER-
ATED BY STRAIGHT LINES AND WITH A HIGH
RATE OF THICKNESS TAPER AT THE ROOT.
D. Peckham. May 1957. 17p. diagrs. RAE
Tech.Note Aero 2451. (Ask for N-50134*)

This note describes a way in which wings can be
designed to have a high rate of thickness taper at the
root, while still maintaining a surface shape gen-
erated by straight lines. The method can be most
successfully applied to wings of parabolic arc sec-
tion straight-tapered plan form, in which case
there is no change in airfoil section shape across
the span. Other plan-form shapes, and wing root
airfoil section shapes, result in a variation of air-
foil section shape across the span.


N-50143*

Royal Aircraft Establishment (Gt. Brit.)
HOUSINGS FOR STANDARD AIRBORNE ELECTRIC
MOTORS. Z. R. S. Ratajski. March 1957. 25p.
diagrs. RAE Tech.Note EL. 135.
(Ask for N-50143*)

This Technical Note discusses the present position
on the standardization of housings for small d-c and
a-c airborne motors and gives details of the modern,
clamp-mounted cylindrical housings already used for
a number of servo-components and motors. Refer-
ence is made to American Standards and proposals,
including those of the Society of Automotive
Engineers, Inc. (U.S.A.), for commercial airborne
motors. It is recommended that the clamp-mounted
cylindrical housing should be adopted as standard
for the whole range of small airborne d-c and a-c
motors.



N-50144*

Royal Aircraft Establishment (Gt. Brit.)
BRIEF LANDING TRIALS USING THE FLUSH
LIGHTING PATTERN IN THE RUNWAY AT THE
ROYAL NETHERLANDS AIR FORCE STATION AT
SOESTERBERG. E. N. Hooton. March 1957.
20p. diagrs. RAE Tech.Note. BL.45.
(Ask for N-50144*)

At the Royal Netherlands Air Force airfield at
Soesterberg a center line and cross bar approach
lighting pattern is installed within the runway flush
with the runway surface. Brief flight tests have
been made using Varsity and Devon aircraft to exam-
ine the visual guidance provided by this approach
pattern during final flare and landing. Although, in
detail, far from ideal for such a purpose, the pro-
vision of a pattern within the runway was considered
to be a great improvement on runway edge lighting
by three of the four pilots who took part in the tests.
The results in general confirm ground simulator
tests and indicate certain elements essential to any
visual guidance system for landing.


,AVAILABLE ON LOAN ONLY










BRITISH REPORTS


N-50145*

Royal Aircraft Establishment (Gt. Brit.)
FATIGUE LOADINGS IN FLIGHT-LOADS IN THE
TAILPLANE OF A COMET I. Anne Burns. March
1957. 19p. diagrs., photo., tabs. RAE Tech.Note
Structures 222. (Ask for N-50145*)

Data are presented on the number of load cycles of
various magnitudes occurring in the tailplane of a
Comet IA during normal ground and flight conditions.
The conditions include flight in turbulence, take-off,
landing, taxiing, and ground running of the engine.
The relative importance of the loads in the different
conditions is illustrated by reference to the loads in
a typical flight.



MISCELLANEOUS


Errata on NACA Rept. 1295, by Rogallo, Yaggy,
and McCloud, In. 1956.


Errata on NACA RM L55L23b, by Hill, Adamson,
Foland, and Bressette. March 26, 1957.


Errata on NACA TN 3176, by Baldwin, Jr.,
Turner, and Knechtel. May 1954.


Errata on NACA TN 3433, by Cunningham.
May 1955.



UNPUBLISHED PAPERS



N-55837*

Battelle Memorial Institute. THE EFFECT OF
VARIATIONS IN MELTING AND CASTING PROCE-
DURE ON THE STRUCTURE AND HIGH-
TEMPERATURE PROPERTIES OF CAST X-40
ALLOY. E. E. Fletcher, H. J. Hucek, A. R.
Elsea, and G. K. Manning. January 16, 1957.
(iii), 59p. diagrs., photos., tabs. Battelle
Memorial Inst. (Ask for N-55837*)

Melting and pouring atmospheres, such as vacuum,
which lowered the nitrogen content of X-40 alloy
below some minimum value (approximately 0.05 to
0.07 percent) markedly reduced the high-temperature
strength of the alloy. The loss of other elements
during melting had a relatively minor effect on the
properties. Variations in pouring temperature over
the range of 26500 to 28500 F had little effect on the
properties. Of the three mold preheat temperatures
studied (3000, 16000, and 19000 F), castings made
in 19000 F molds yielded the best and most consist-
ent high-temperature properties. No consistent re-
lationship between microstructure and high-
temperature properties was observed; however,


NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO 122

there were indications that castings with an inter-
mediate grain size were stronger than those with
either extremely coarse or fine gram sizes. The
high-temperature bend test was found to be unsuit-
able as a screening test for stress-rupture proper-
ties, but it appeared to be useful for evaluating de-
formation or creep characteristics.


N-55838'

Battelle Memorial Institute. THE EFFECTS OF
VARIATIONS IN NITROGEN AND MANGANESE CON-
TENT ON THE STRUCTURE AND HIGH-
TEMPERATURE PROPERTIES OF CAST X-40
ALLOY. E. E. Fletcher, B. J. Hucek, A. R.
Elsea, andG. K. Manning. June 14, 1957. 32p.
diagrs., photos., tabs. Battelle Memorial Inst.
(Ask for N-55838*)



N-55839*

Polytechnic Institute of Brooklyn. ON SOME EX-
ACT SOLUTIONS OF LAMINAR MIXING IN THE
PRESENCE OF AXIAL PRESSURE GRADIENTS.
Luigi G. Napolitano. December 1955. 22p.
Polytechnic Inst. of Brooklyn. PIBAL Report 302.
(Ask for N-55839*)

The well-known power and exponential velocity dis-
tributions in the Stewartson plane are shown to still
yield similar solutions in the case of compressible
mixing provided an additional requirement, concern-
ing the Mach number distributions 6f the two
streams, is satisfied. Analysis of the correspond-
ing incompressible problem shows that no similar
solutions are possible except for the simplest case
of zero pressure gradient. For the mixing of two
uniform streams the ordinary differential equation
giving the similar solutions is shown to reduce to
the Blasius equation in both the compressible and the
incompressible cases. Consequently with this ap-
proach the time-consuming iteration process which
arises in the previously given methods of analysis is
eliminated. The possibility of a rapid investigation
of temperature effect on wake characteristics is
stressed.



N-55840*

Polytechnic Institute of Brooklyn. EXPERIMENTAL
INVESTIGATION OF THE MIXING OF TWO HOMO-
GENEOUS STREAMS IN CHANNELS OF GIVEN GE-
OMETRY. Luigi G. Napolitano and Marian Visich,
Jr. January 1957. 88p. diagrs., photos.
Polytechnic Inst. of Brooklyn. PIBAL Report 321.
(Ask for N-55840*)

Four series of tests were performed in a constant
area channel, each series being characterized by a
value of the Mach number of stream A and a value
of the height of stream B at the origin of mixing.
M = 1.6 and 2.5; the heights of stream B at the ori-
gin of mixing were equal to 1 and 2 inches. Two
additional series of tests were performed n a vari-
able area channel at M = 2.5 for the two previously
mentioned heights of stream B. Overall character-
istics of mixing process are presented in terms of
static pressure distributions along the channel, Mach


* AVAILABLE ON LOAN ONLY





NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122 7


UNPUBLISHED PAPERS


number profiles at the discharge end, static pres-
sures at the origin of mixing, and schlieren photo-
graphs of relevant flow patterns. Effects of several
parameters are discussed and a correlation of some
Sof the experimental data is presented. Applications
of the results to some practical problems such as
Supersonic ejector design and base bleed are
outlined.




DECLASSIFIED
NACA RESEARCH REPORTS

This is authority for those possessing copies
of the following research reports to declassify
them from Confidential. Items marked (#)
are UNAVAILABLE since they contain obso-
lete or proprietary information. Effective
date of declassification November 8, 1957.


#NACA MR L6G12a, by Gilman and Harman
(5321 Army-Navy UPA-11/1)


NACA RM A53K20

INVESTIGATION OF A TRAILING-EDGE PADDLE-
CONTROL SURFACE ON A TRIANGULAR WING OF
ASPECT RATIO 2 AT SUBSONIC AND SUPERSONIC
SPEEDS. Louis H. Ball. February 1954. 18p.
diagrs., tab. (NACA RM A53K20)

This report presents the results of an experimental
investigation of external airfoils known as paddle-
control surfaces as the longitudinal control device on
a triangular wing of aspect ratio 2. The lift, drag,
pitching moment, and hinge moment were obtained
for Mach numbers of 0.60, 0.80, 0.90, 1.20, 1.30,
1.50, 1.70, and 1.90 at a constant Reynolds number
of 3.0 x 106, for angles of attack from about -40 to
180 and for paddle-control deflections from approxi-
mately 40 to -160.


NACA RM A54K09

LONGITUDINAL STABILITY CHARACTERISTICS
AT MACH NUMBERS UP TO 0.92 OF A WING-
BODY-TAIL COMBINATION HAVING A WING WITH
450 OF SWEEPBACK AND A TAIL IN VARIOUS
VERTICAL POSITIONS. Jack D. Stephenson,
Angelo Bandettini, and Ralph Selan. January 1955.
64p. diagrs., photos., tabs. (NACA RM A54K09)

Data are presented showing lift, drag, and static
longitudinal staOnlily characteristics of a wing-body-
tail combination having a 450 sweptback wing with an
aspect ratio of 5.5. Lowering the horizontal tail to
the wing chord plane extended resulted in significant
improvements in the stability of the model both with
an unmodified wing and with fences or chord exten-
sions on the wing.


NACA RM E9K29

ANALYTICAL AND EXPERIMENTAL INVESTIGA-
TION OF THRUST AUGMENTATION OF AXIAL-
AND CENTRIFUGAL-COMPRESSOR TURBOJET
ENGINES BY INJECTION OF WATER AND ALCO-
HOL IN COMBUSTION CHAMBERS. David S.
Gabriel, Harry W. Dowman, and William L. Jones.
April 13, 1950. 43p. diagrs., photo.
(NACA RM E9K29)

Thrust augmentation by injection of water and water-
alcohol mixtures into combustion chambers was
investigated at sea-level zero-ram conditions on
axial- and centrifugal-flow turbojet engines. Thrust
augmentation, compressor characteristics, fuel
flow, and turbine-outlet-temperature distributions
are presented for various injection rates and water-
alcohol mixtures. A method of computing thrust
augmentation produced by injection into the com-
bustion chambers is presented. Computed
augmented-thrust ratios were within 1 percent of
the experimental results. Water injection into the
centrifugal-flow engine produced little thrust
augmentation.


#NACA RM SE50J12, by Cook and Butze


NACA RM E50K22

EXPERIMENTAL INVESTIGATION OF TAIL-PIPE-
BURNER DESIGN VARIABLES. W. A. Fleming,
E. William Conrad, and A. W. Young. March 5,
1951. 75p. diagrs., photos., tab.
(NACA RM E50K22)

The results of several experimental tail-pipe-burner
investigations at the NACA Lewis Laboratory are
summarized to indicate the effects of tail-pipe-
burner design variables on performance and operat-
ing characteristics. Most of the configurations
were operated over wide ranges of altitudes and
flight Mach numbers. The data indicate the effect
of changes in principal design variables, such as
flame-holder type, burner length, burner diameter,
and fuel distribution on performance, and indicate
the features of a tail-pipe burner that will operate
with high combustion efficiency up to an altitude of
approximately 50,000 feet.


NACA RM E51A25

INVESTIGATION OF ALTITUDE IGNITION, ACCEL-
ERATION AND STEADY-STATE OPERATION WITH
SINGLE COMBUSTOR OF J47 TURBOJET ENGINE.
William P. Cook and Helmut F. Butze. March 5,
1951. 35p. diagrs., photo., tab.
(NACA RM E51A25)

Satisfactory ignition was obtained with a J47 single
combustor up to and including an attitude of 40,000
feet at conditions simulating equilibrium i ndmilling
of the engine at flight speed of 400 miles per hour.
At 30,000 feet, excess temperature rise available
for combustion was limited by the ability of the
combustor to provide temperature rise, whereas at
high engine speeds the maximum allowable turbine-
inlet temperature became the restricting factor.


*AVAILABLE ON LOAN ONLY.
# UNAVAILABLE; contains proprietary or obsolete information.









DECLASSIFIED NACA RESEARCH REPORTS


Altitude operational limits increased from about
51,500 feet at 55-percent rated speed to about 64,500
feet at 85-percent rated speed. Combustion effi-
ciencies varied from 39.0 to 92.6 percent.



INACA RM E51K06, by Prince and Wintler



NACA RM E51K20

ANALYTICAL AND EXPERIMENTAL INVESTIGA-
TION OF INLET-ENGINE MATCHING FOR
TURBOJET-POWERED AIRCRAFT AT MACH NUM-
BERS UP TO 2.0. Carl F. Schueller and Fred T.
Esenwein. February 1952. 31p. diagrs., photos.
(NACA RM E51K20)

An analysis of inlet-turbojet-engine matching for
a range of Mach numbers up to 2.0 indicates large
performance penalties when fixed-geometry inlets
are used. Use of variable-geometry inlets, how-
ever, nearly eliminates these penalties. The
analysis was confirmed experimentally by investi-
gating at Mach numbers of 0, 0.63, and 1.5 to 2.0
two single oblique-shock-type inlets of different
compression-ramp angles, which simulated a
variable-geometry configuration. The experimental
investigation indicated that total-pressure recov-
eries comparable with those attainable with well
designed nose inlets were obtained with the side
inlets when all the boundary layer ahead of the
inlets was removed. Serious drag penalties
resulted at a Mach number of 2.0 from the use of
blunt-cowl leading edges. However, sharp-lip
inlets produced large losses in thrust for the take-
off condition. These thrust penalties which are
associated with the low-speed operation of the
sharp-lip inlet designs can probably be avoided
without impairing the supersonic performance of the
inlet by the use of auxiliary inlets or blow-in doors.



NACA RM E51K29

LOITERING AND RANGE PERFORMANCE OF
TURBOJET-POWERED AIRCRAFT DETERMINED
BY OFF-DESIGN ENGINE CYCLE ANALYSIS.
Stanley L. Koutz and Reece V. Hensley.
February 1952. 45p. diagrs., tab.
(NACA RM E51K29)

The loitering and range performance of airplanes
equipped with several different turbojet engines was
analytically investigated by applying the results of
off-design cycle analyses to specific airplane char-
acteristics. The method of off-design cycle analy-
sis is presented herein and is verified by a check
with experimental data. For all engines considered,
the loitering and the range fuel flows obtained with
rated tail-pipe nozzle area, variable engine speed
operation were within 2 or 3 percenE of the optimum
fuel flow obtainable with any method of engine opera-
tion. The optimum loitering altitude generally
occurred between approximately 25,000 and 35,000
feet with corresponding optimum flight Mach num-
bers of 0.4 to 0.65. In general, the optimum range


*AVAILABLE ON LOAN ONLY


NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122


fuel flows occurred at 3000 to 5000 feet higher alti-
tude and at approximately 0.15 higher flight Mach
numbers than the optimum loitering fuel flow.



NACA RM E52K04

INTERNAL PERFORMANCE OF SEVERAL TYPES
OF JET-EXIT CONFIGURATIONS FOR SUPERSONIC
TURBOJET AIRCRAFT. William A. Fleming.
January 1953. 28p. diagrs. (NACA RM E52K04)

Internal performance characteristics of a convergent
nozzle, fixed convergent-divergent nozzles, an
adjustable plug-type convergent-divergent nozzle,
and jet ejectors were investigated over a wide range
of pressure ratios. These data are summarized to
provide an overall picture of jet-exit performance
and thereby serve as an aid in selecting the most
suitable jet-exit configurations for supersonic
aircraft.



NACA RM E52K14

FORCE AND PRESSURE RECOVERY CHARACTER-
ISTICS AT SUPERSONIC SPEEDS OF A CONICAL
SPIKE INLET WITH BYPASSES DISCHARGING [N
AN AXIAL DIRECTION. J. L. Allen and Andrew
Beke. January 1953. 27p. diagrs., photos., tab.
(NACA RM E52K14)

Force and pressure-recovery characteristics of a
nacelle-type conical-spike inlet with two fixed-area
bypasses are presented for flight Mach numbers of
1.6, 1.8, and 2.0 and for angles of attack from 00 to
90. The inlet was designed to attain a mass-flow
ratio of unity at a flight Mach number of 2.0. The
horizontally opposed bypasses were 6 inlet diameters
aft of the inlet entrance. At a flight Mach number of
2.0, the discharge of 23 percent of the critical mass
flow of the inlet by means of bypasses (thus retain-
ing critical inlet flow) increased the drag only one-
fifth of the additive drag that would result if the
same amount of air were spilled behind on inlet
normal shock. The total-pressure recovery of the
diffuser was not significantly reduced. Similar
results were obtained at other flight Mach numbers.


NACA RM E52K17

PERFORMANCE OF DOUBLE-SHROUD EJECTOR
CONFIGURATION WITH PRIMARY PRESSURE
RATIOS FROM 1.0 TO 10. Donald P. Hollister and
William K. Greathouse. February 1953. 34p.
diagrs., tabs. (NACA RM E52K17)

A brief investigation was made to determine the per-
formance characteristics of a double-shroud
cooling-air ejector configuration. Two convergent
primary nozzles were used to simulate a specific
manufacturer's iris-type variable-area nozzle in
the open and closed positions. The investigation
comprised four phases: (1) obtained performance
with no secondary or tertiary air flow (cooling-air
passages blocked), (2) determining the tendency for
backflow to occur in either cooling-air passage,
(3) determining the sensitivity of flow in one passage
to that in the other, and (4) obtaining pumping and
thrust characteristics with secondary and tertiary


# UNAVAILABLE; contains proprietary or obsolete information.






NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122 9


DECLASSIFIED NACA RESEARCH REPORTS


air flow. The experimental results showed that
the performance with the cooling-air passages
blocked was typical of that for single ejectors
having diameter and spacing ratios similar to those
used in this investigation. There was a tendency
S for backflow to occur, but the magnitude of such
flow was relatively small. The weight flow of each
ejector was shown to be essentially independent of
the other, and the closed primary-nozzle config-
uration was found to be generally capable of pump-
ing more cooling air than the open primary-nozzle
configuration. Gains in gross thrust were observed
for both configurations, with losses occurring only
at low secondary and tertiary pressure ratios.


NACA RM E52K18

INVESTIGATION AT SUPERSONIC SPEEDS OF AN
INLET EMPLOYING CONICAL FLOW SEPARATION
FROM A PROBE AHEAD OF A BLUNT BODY.
Donald P. Hearth and Gerald C. Gorton. January
1953. 32p. diagrs., photos., tab.
(NACA RM E52K18)

An experimental investigation was conducted on an
inlet employing conical flow separation from a probe
extending upstream from a hemispherical-nosed
centerbody. Data were obtained at free-stream
Mach numbers from 1.6 to 2.0 and angles of attack
from 00 to 90. Pressure-recovery and drag char-
acteristics for the inlet were very nearly comparable
with those for a conical-spike inlet at zero angle of
attack and design Mach number of 2.0, but compared
less favorably at Mach numbers below 2.0. A large
reduction in pressure recovery and mass flow was
obtained at angle of attack. However, an
investigation on the use of probes offset from the
inlet center line indicated that the angle-of-attack
performance could be appreciably improved if the
probe were alined with the stream direction.


NACA RM E52K26

ANALYSIS OF SEVERAL METHODS OF PUMPING
COOLING AIR FOR TURBOJET-ENGINE AFTER-
BURNERS. John C. Samuels and Herbert Yanowitz.
February 1953. 54p. diagrs. (NACA RM E52K26)

Several methods of pumping air to an annular cooling
passage surrounding a typical axial-flow turbojet-
engine afterburner were evaluated and compared on
the basis of thrust and specific fuel consumption of
the systems. Each system was analyzed over a
range of afterburner-wall temperatures, flight Mach
numbers, and exhaust-gas temperatures at sea level
and 35,000 feet. Ram pressure recovery, boundary-
layer pressure recovery, and the engine-jet actuated
ejector appear to be satisfactory systems at high
Mach numbers. Cooling with compressor-exit air
bleed was found to be unsatisfactory, but the use of
compressor-exit bleed air as the primary fluid in a
high-performance ejector was satisfactory. The use
of an auxiliary compressor driven from the engine
shaft increased the thrust and decreased the specific
fuel consumption of the engine for many of the condi-
tions investigated.


NACA RM E53K03

ANALYSIS OF PERFORMANCE OF FOUR
SYMMETRICAL-DIAGRAM-TYPE SUBSONIC
INLET-STAGE AXIAL-FLOW COMPRESSORS.
Robert J. Jackson. January 1954. 72p. diagrs.,
photos. (NACA RM E53K03)

The main results of an experimental investigation of
four axial-flow single-stage compressors having a
hub-tip radius ratio of 0.5 and designed for symmet-
rical velocity diagrams at all radii were as follows:
Because of the action of radial equilibrium, the high-
pressure-ratio and low-weight-flow type of design
incurred high tip-region blade loadings and, conse-
quently, high losses. Therefore, the most suitable
use for the symmetrical-velocity-diagram type of
design is for higher specific weight flow and (for a
given Mach number limit) lower pressure ratio de-
signs. A diffusion factor correlation was obtained,
for a given solidity, among compressor and
cascade-predicted deviation angles, and among
compressor losses for the range of diffusion factor
exceeding approximately 0.5. The available data
indicated that the assumption of simplified radial
equilibrium was a valid one. In order to predict
the radial variation of air velocities, it was neces-
sary to consider the radial variation of entropy,
which was especially significant after the stators.



NACA RM E54I27a

PRELIMINARY INVESTIGATION OF THE STRENGTH
AND ENDURANCE OF PLASTIC-IMPREGNATED
FIBERGLASS COMPRESSOR BLADES. Donald F.
Johnson and Andre J. Meyer, Jr. January 1955.
21p. diagrs., photos. (NACA RM E54I27a)

A complete set of third-stage compressor rotor
blades fabricated from laminated fiberglass cloth
and bonded with a phenolic resin were operated in
a conventional engine for 105 hours. All blades
were in perfect condition upon disassembly. Also,
a number of laboratory tests were conducted on
additional blades to determine moduli of elasticity
and rigidity, strength in both tension and bending,
internal damping, and the fatigue limit of the blade
material.



NACA RM E54K02

PRELIMINARY INVESTIGATION OF A TECHNIQUE
OF PRODUCING A HEATED CORE IN A SUPER-
SONIC WIND-TUNNEL STREAM. Morris D.
Rousso and Milton A. Beheim. February 1955.
22p. photos., diagrs. (NACA RM E54K02)

An investigation conducted at Mach numbers 1.9 and
3.0 has shown that a central core of air of high
stagnation temperature can be produced in the test
section of a supersonic wind tunnel. Air heated by
combustion was injected into the tunnel in a stream-
wise direction near the tunnel throat. The effects
of core-nozzle location and dimensions, core to
main air stream total-pressure ratio, and core
stagnation temperature on the size and shape of
temperature contours and on Mach number profiles
in the test section were determined.


*AVAILABLE ON LOAN ONLY









DECLASSIFIED NACA RESEARCH REPORTS


NACA RM L8K04a

FLIGHT TESTS AT TRANSONIC AND SUPERSONIC
SPEEDS OF AN AIRPLANE-LIKE CONFIGURATION
WITH THIN STRAIGHT SHARP-EDGE WINGS AND
TAIL SURFACES. Clarence L. Gillis and Jesse L.
Mitchell. January 5, 1949. 37p. diagrs., photos.,
tab. (NACA RM L8K04a)

Rocket-powered models of a configuration having
thin, straight, sharp-edge wings and tail surfaces
were flight-tested at transonic and low supersonic
speeds. Rather large longitudinal trim changes
occurred. Decreased longitudinal control effective-
ness was evident at supersonic speeds. The static
directional stability appears adequate but a snaking
oscillation generally existed at supercritical Mach
numbers.


NACA RM L8K05

DRAG MEASUREMENTS IN FLIGHT ON THE
10-PERCENT-THICK AND 8-PERCENT-THICK
WING X-l AIRPLANES. John J. Gardner.
November 19, 1948. 17p. diagrs., photo.
(NACA RM L8K05)

Contains results of drag measurements of X-l air-
planes with 8-percent-thick wing, 6-percent-thick
tail and 10-percent-thick wing, 8-percent-thick tail
in Mach number range from 0.7 to 1.3. Brief com-
parison made between airplane and wing section lift-
drag ratios.


NACA RM L9D08

ESTIMATED TRANSONIC FLYING QUALITIES OF A
TAILLESS AIRPLANE BASED ON A MODEL INVES-
TIGATION. Charles J. Donlan and Richard E. Kuhn.
June 8, 1949. 63p. diagrs., photos., tabs.
(NACA RM L9D08)

An analysis of the estimated flying qualities of a tail-
less airplane with the wing quarter-chord line swept
back 350, an aspect ratio of 3, and taper ratio of 0.6
in the Mach number range from 0.40 to 0.91 has been
made. The longitudinal stability and control were
investigated and estimates made of the period and
damping of the short-period longitudinal and lateral
oscillations. Also included are the wind-tunnel test
data on which the analysis is based. Data include
longitudinal and lateral stability and control-force
data, effect of speed brakes, canopies, fins, and
wing-alone tests. Tuft studies of flow on wings, fins,
and behind speed brakes are included throughout the
Mach number range.


NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122

NACA RM L9J07a

INVESTIGATION OF THE DYNAMIC LATERAL
STABILITY AND CONTROL CHARACTERISTICS OF
A MODEL OF A FIGHTER AIRPLANE WITHOUT A
HORIZONTAL TAIL AND EQUIPPED WITH EITHER
SINGLE OR TWIN VERTICAL TAILS. John W.
Draper and Robert W. Rose. November 15, 1949.
20p. diagrs., photos., tab. (NACA RM L9J07a)

An investigation has been conducted in the Langley
free-flight tunnel to compare the lateral stability and
control characteristics of a model of a fighter air-
plane without a horizontal tail and equipped either
with a single vertical tail mounted on the fuselage or
with twin vertical tails of the same tail volume
mounted on the wing. The investigation included
force and flight tests of the model with both tail
configurations.






#NACA RM SL50B23, by Michal






#NACA RM SL50H23a, by Ulmann and Lord






NACA RM L50K03

INVESTIGATION OF THE AERODYNAMIC EFFECTS
OF AN EXTERNAL STORE IN COMBINATION WITH
600 DELTA AND LOW-ASPECT-RATIO TAPERED
WINGS AT A MACH NUMBER OF 1.9. EUery B.
May, Jr. January 9, 1951. 46p. diagrs., photos.
(NACA RM L50K03)

A wind-tunnel investigation was made of an external
store of fineness ratio 8.6 which was located on the
chord plane of several wings in the Langley 9- by
12-inch supersonic blowdown tunnel. The work was
carried out at a Mach number of 1.9 on wings having
600 leading-edge sweepback and unswept trailing
edges with taper ratios of 0 and 0.28, and on an un-
swept wing having a taper ratio of 0.625. Investiga-
tion was made to determine the aerodynamic effects
of store position, the loading breakdown between the
store and the inner and outer wing panel of one wing,
and the effects of the store on control characteristics
The test Reynolds numbers ranged between 2.3 x 106
and 4.0 x 106.


* AVAILABLE ON LOAN ONLY.
# UNAVAILABLE; contains proprietary or obsolete information.





NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122 11


DECLASSIFIED NACA RESEARCH REPORTS


NACA RM L50K16

WING-ON AND WING-OFF LONGITUDINAL CHAR-
ACTERISTICS OF AN AIRPLANE CONFIGURATION
S HAVING A THIN UNSWEPT TAPERED WING OF
ASPECT RATIO 3, AS OBTAINED FROM ROCKET-
PROPELLED MODELS AT MACH NUMBERS FROM
0.8 TO 1 4. Clarence L. Gillis and A. James Vitale.
March 14, 1951. 52p. diagrs., photos., tabs.
(NACA RM L50K16)

Flight tests were conducted on three rocket-
propelled models of an airplane configuration, two
models having thin unswept tapered wings of aspect
ratio 3 and hexagonal airfoil sections, and the third
having no wing. The two winged models had wings
of different stiffness characteristics. Aerodynamic
derivatives defining the static and dynamic longitu-
dinal stability, control, trim, and drag character-
istics of the configuration were obtained over a Mach
number range from 0.8 to 1.4. For some of the
aerodynamic derivatives the separate effects of the
wing, tail, and fuselage were determined. Some
buffeting and maximum lift information was obtained
at high subsonic speeds.



NACA RM L51K05

TOTAL-PRESSURE RECOVERY OF A CIRCULAR
UNDERSLUNG INLET WITH THREE DIFFERENT
NOSE SHAPES AT A MACH NUMBER OF 1.42.
Charles F. Merlet and Howard S. Carter. February
1952. 37p. diagrs., photos. (NACA RM L51K05)

Total-pressure recoveries at the inlet and after dif-
fusion and shadowgraphs are presented for a circular
underslung inlet located well forward on a body of
revolution. Three nose shapes, varying in blunt-
ness, were tested at three angles of attack and one
angle of yaw in a free-air jet at a Mach number of
1.42, and over a range of mass-flow ratios from 0.3
to 0.9.


NACA RM L51K06

LONGITUDINAL STABILITY AND DRAG CHARAC-
TERISTICS AT MACH NUMBERS FROM 0.75 TO 1.5
OF AN AIRPLANE CONFIGURATION HAVING A 600
SWEPT WING OF ASPECT RATIO 2.24 AS OB-
TAINED FROM ROCKET-PROPELLED MODELS.
A. James Vitale, John C. McFall, Jr., and John D.
Morrow. April 1952. 43p. diagrs., photos., tabs.
(NACA RM L51K06)

Flight tests were conducted on a rocket-propelled
airplane configuration model and on a drag model,
each having a 600 swept wing of aspect ratio 2.24
with different fuselages. The longitudinal stability,
control, and drag characteristics of the airplane
configuration were determined over a Mach number
range of 0.75 to 1.50. Wing-plus-interference mini-
mum drag was obtained from the drag model over a
Mach number range of 0.90 to 1.50. For some of the
aerodynamic derivatives the separate effects of the
wing, tail, and fuselage were obtained.


NACA RM L51K20

SUMMARY OF SOME EFFECTIVE AERODYNAMIC
TWISTING-MOMENT COEFFICIENTS OF VARIOUS
WING-CONTROL CONFIGURATIONS AT MACH
NUMBERS FROM 0.6 TO 1.7 AS DETERMINED
FROM ROCKET-POWERED MODELS. H. Kurt
Strass. January 1952. 22p. diagrs., photo.,
2 tabs. (NACA RM L51K20)

This paper presents a summary of some effective
aerodynamic twisting-moment coefficients of various
wing-control configurations at Mach numbers from
0.6 to 1.7 as determined by the use of rocket pro-
pelled test vehicles which indicated that, within the
framework of the necessary assumptions, the value
of the effective twisting-moment coefficient de-
creased as the sweepback of the aileron hinge axis
is increased. Large changes in the value of the ef-
fective twisting-moment coefficient were obtained in
the Mach number region from M 0.8 to M 1.2
with changes in aileron span and location upon the
same wing plan form. Above M 1. 2, the values
tended to agree more closely. This factor limits the
use of these data to wing-control configurations sim-
ilar to those tested. Comparative tests of an out-
board 0.3-span, 0.25-chord aileron, and a midspan
spoiler of approximately the same span length indi-
cate that the twisting moment of the spoiler is about
one-third that of the aileron for equal values of roll-
ing effectiveness.


NACA RM L52K17

INVESTIGATION AT TRANSONIC SPEEDS OF A
FORWARD-LOCATED UNDERSLUNG AIR INLET
ON A BODY OF REVOLUTION. P. Kenneth
Pierpont and John A. Braden. January 1953.
109p. diagrs., photos., tabs. (NACA RM L52K17)

Results are given for an experimental investigation
of a forward-located underslung scoop mounted on a
basic body at Mach numbers from 0.6 to 1.1 and for
angles of attack from 00 to 100. The inlet area was
17.3 percent of the body frontal area. For mass-
flow ratios from 0.3 to within 5 percent of the maxi-
mum, total pressure recovery in the diffuser in
excess of 97 percent was realized at all Mach num-
bers and angles of attack. External drag of the
scoop model at high mass-flow condition was approx-
imately equal to that of the basic body.


NACA RM L52K20a

EFFECTS OF SIZE OF EXTERNAL STORES ON THE
AERODYNAMIC CHARACTERISTICS OF AN UN-
SWEPT AND A 450 SWEPTBACK WING OF ASPECT
RATIO 4 AND A 600 DELTA WING AT MACH NUM-
BERS OF 1.41, 1.62, AND 1.96. Carl R. Jacobsen.
January 1953. 55p. diagrs., photos., tab.
(NACA RM L52K20a)

An investigation has been made in the Langley 9- by
12-inch supersonic blowdown tunnel to determine the
effects of external stores of various sizes on the lift,
drag, and pitching-moment characteristics of three
wings; an unswept and a 450 sweptback wing having
aspect ratios of 4 and taper ratios of 0.6, and a 600
delta wing at Mach numbers of 1.41, 1.62, and 1.96.


'AVAILABLE ON LOAN ONLY.





NACA RESEARCH ABSTRACTS
12 AND RECLASSIFICATION NOTICE NO. 122


DECLASSIFIED NACA RESEARCH REPORTS



The size of a Douglas Aircraft Company, Inc., store
having a fineness ratio of 8.58 was systematically
varied at the 80-percent-semispan station of the
unswept and sweptback wini and at the 60-percent-
semispan station of the delta wing. Several stores
were also tested at various other store locations on
the outer 60 percent of each of the wing semispans.
For wing areas of 500, 600, and 750 square feet for
an unswept, a sveptback, and a delta wing, respec-
tively, the store sizes covered in the investigation
provide data for stores ranging in size from a 200-
pound bomb to a large jet engine nacelle.



NACA RM L52K21

EFFECT OF LEADING-EDGE CHORD-EXTENSIONS
ON SUBSONIC AND TRANSONIC AERODYNAMIC
CHARACTERISTICS OF THREE MODELS HAVING
450 SWEPTBACK WINGS OF ASPECT RATIO 4.
Kenneth W. Goodson and Albert G. Few, Jr.
January 1953. 31p. diagrs., photos., tab.
(NACA RM L52K21)

This paper presents the effects of wing leading-edge
chord-extensions on the subsonic and transonic high-
lift pitching moment and the performance character-
istics of three models having wings of 450 sweep-
back, aspect ratio 4, taper ratio 0.30, and NACA
65A006 airfoil sections parallel to the free stream.
The test Reynolds numbers varied from about
0.6 x 106 to 3.25 x 106, depending upon the wing and
Mach number. Some tests were made with a full-
chord fence installed.



NACA RM L52K28

HINGE-MOMENT CHARACTERISTICS FOR SEVER-
AL TIP CONTROLS ON A 600 SWEPTBACK DELTA
WING AT MACH NUMBER 1.61. K. R. Czarnecki
and Douglas R. Lord. January 1953. 31p. diagrs.,
photos. (NACA RM L52K28)

An investigation has been made at a Mach number of
1.61 and a Reynolds number of 4.2 x 106 to deter-
mine the hinge-moment characteristics of a group
of tip controls on a 600 sweptback delta wing. The
control configurations varied in plan form and in
amount of aerodynamic balance. Tests were made
over an angle-of-attack range of 0 to 150 and a flap
deflection range of -300 to 300. A comparison of
the experimental results with theory is included.



#NACA RM SL53H11, by Falanga


NACA RM L53K03

EFFECT OF LARGE DEFLECTIONS OF A CANARD
CONTROL AND DEFLECTIONS OF A WING -TIP
CONTROL ON THE STATIC-STABILITY AND
INDUCED-ROLL CHARACTERISTICS OF A CRUCI-
FORM CANARD MISSILE AT A MACH NUMBER OF
2.01. M. Leroy Spearman. December 1953. 20p
diagrs., tabs. (NACA RM L53K03)

An investigation has been conducted in the Langley
4- by 4-foot supersonic pressure tunnel at a Mach
number of 2.01 to determine the effect of large de-
flections of a canard and deflections of a wing-tip
control on the static stability and induced-roll char-
acteristics of a cruciform-wing canard-type missile.
The missile had a body fineness ratio of 15.7. The
tests covered an angle-of-attack range from 0 up to
about 270, a canard deflection range up to 300, and a
wing-tip control deflection up to 200



NACA RM L53K11

EFFECTS OF A SERIES OF INBOARD PLAN-FORM
MODIFICATIONS ON THE LONGITUDINAL CHAR-
ACTERISTICS OF TWO UNSWEPT WINGS OF AS-
PECT RATIO 3.5, TAPER RATIO 0.2, AND DIF-
FERENT THICKNESS DISTRIBUTIONS AT MACH
NUMBERS OF 1.61 AND 2.01. John R. Sevier, Jr.
February 1954. 43p. diagrs., photos.
(NACA RM L53K11)

Tests of a series of inboard plan-form modifications
of two unswept wings of aspect ratio 3.5 and taper
ratio 0.2 were made at the Langley 4- by 4-foot
supersonic pressure tunnel at Mach numbers of 1.61
and 2.01. Results indicated that, by properly modi-
fying the wing thickness and plan form, significant
increases in wing volume can be attained with little
or no penalties in drag and actual increases in maxi-
mum lift-drag ratio.



NACA RM L53K27

WIND-TUNNEL INVESTIGATIONS AT LOW AND
TRANSONIC SPEEDS OF THE FEASIBILITY OF
SELF-ACTUATING SPOILERS AS A LATERAL-
CONTROL DEVICE FOR A MISSILE. Harleth G.
Wiley and William C. Hayes, Jr. January 1954.
24p. diagrs., tab. (NACA RM L53K27)

Results are presented of wind-tunnel investigations
of the lateral-control effectiveness of sell-actuating
spoilers on a thin 600 delta wing up to an angle of
attack of about 200 at low speeds and of the aero-
dynamic moments acting on an isolated self-
actuating spoiler through an angle-of-rotation range
of 1800 at transonic speeds. Presented also are
the results of brief tests of spoiler actuating times
at low speeds.


*AVAILABLE ON LOAN ONLY
# UNAVAILABLE; contains proprietary or obsolete information.






NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO.122


DECLASSIFIED NACA RESEARCH REPORTS



NACA RM L54K02

A LIMITED FLIGHT INVESTIGATION'OF THE
EFFECT OF DYNAMIC VIBRATION ABSORBERS
ON THE RESPONSE OF AN AIRPLANE STRUCTURE
DURING BUFFETING. Jim Rogers Thompson and
John E. Yeates, Jr. January 1955. 29p. diagrs.,
photos. (NACA RM L54K02)

Limited flight measurements have been made in the
high-speed low-lift buffet region of the effect on the
buffeting response of dynamic vibration absorbers
mounted on the wing tips. The absorbers had 70-
pound-moving weights and greatly reduced the
response in the tuned mode. Approximate calcula-
tions indicated that significant reductions could be
produced by an installation weighing considerably
less than that tested. The device shows promise
both as a buffet alleviator and as a research
instrument.



NACA RM L54K11a

EXPERIMENTAL AERODYNAMIC FORCES AND
MOMENTS AT LOW SPEED OF A MISSILE MODEL
DURING SIMULATED LAUNCHING FROM THE
MIDSEMISPAN LOCATION OF A 450 SWEPTBACK
WING-FUSELAGE COMBINATION. William J.
Alford, Jr., H. Norman Silvers, and Thomas J.
King, Jr. February 1955. 36p. diagrs., photo.,
tabs. (NACA RM L54Klla)

An investigation was made at low speed in the
Langley 300-mph 7- by 10-foot tunnel to determine
the aerodynamic forces and moments of a missile
model during simulated launching from the mid-
semispan location of a 450 sweptback wing-fuselage
combination. Various chordwise and vertical
locations were investigaTed with and without a flat-
sided pylon installed. The angle-of-attack range
generally extended from -8 to 28.



NACA RM L54K12

TRANSONIC LONGITUDINAL AERODYNAMIC
EFFECTS OF SWEEPING UP THE REAR OF THE
FUSELAGE OF A ROCKET-PROPELLED AIR-
PLANE MODEL HAVING NO HORIZONTAL TAIL.
James H. Parks. January 1955. 30p. diagrs.,
photo. (NACA RM L54K12)

The transonic longitudinal aerodynamic effects of
sweeping up the rear of the fuselage have been in-
vestigated in free flight by using rocket-propelled
models having no horizontal tails. The models had
aspect-ratio-4 wings with both the wings and ver-
tical tails swept back 450. Sttic and dynamic
longitudinal stability parameters, trim conditions,
drag at trim lift, and local downflow values are
presented.


NACA RM L54K15a

A THEORETICAL INVESTIGATION OF THE EF-
FECT OF AUXILIARY DAMPING ON THE LONGI-
TUDINAL RESPONSE OF A TRANSONIC BOMBER
CONFIGURATION IN FLIGHT THROUGH CON-
TINUOUS TURBULENCE. T. F. Bridgland, Jr.
March 1955. 26p. diagrs., tab.
(NACA RM L54K15a)

A theoretical investigation has been made of the ef-
fects of auxiliary pitch-rate damping on the longi-
tudinal response of a transonic bomber configuration
in low-altitude flight through continuous rough air.
The methods of generalized harmonic analysis are
utilized in obtaining the statistical character of the
airframe responses to a random gust velocity input.
The results of this investigation indicate that, for
the airframe and speed range considered, reductions
in root-mean-square normal acceleration of about
24 percent and reductions in root-mean-square
pitch angle of about 74 percent are possible through
the use of auxiliary pitch-rate damping.


NACA RM L54K30

FREE-FLIGHT INVESTIGATION, INCLUDING
SOME EFFECTS OF WING AEROELASTICITY, OF
THE ROLLING EFFECTIVENESS OF AN ALL-
MOVABLE HORIZONTAL TAIL WITH DIFFEREN-
TIAL INCIDENCE AT MACH NUMBERS FROM 0.6
TO 1.5. Roland D. English. January 1955. 11p.
diagrs., photo. (NACA RM L54K30)

A free-flight investigation has been made to deter-
mine the rolling effectiveness at zero angle of attack
of an all-movable horizontal tail at a constant angle
of differential incidence of 70 per fin mounted behind
a notched delta wing over a Mach number range from
0.6 to 1.5. Two models were tested, one with a
stiff wing and one with a flexible wing. The rolling
effectiveness of both models was of about the same
magnitude at subsonic and supersonic speeds. The
rolling effectiveness of the flexible-wing model was
about 1.6 times that of the stiff-wing model except
in the transonic region.


NACA RM L55A24

FLIGHT AND PREFLIGHT TESTS OF A RAM JET
BURNING MAGNESIUM SLURRY FUEL AND
UTILIZING A SOLID-PROPELLANT GAS GENERA-
TOR FOR FUEL EXPULSION. Walter A. Bartlett,
Jr., and William K. Hagginbothom, Jr. April 1955.
35p. diagrs., photos. (NACA RM L55A24)

Data obtained from the first flight test of a ram jet
utilizing a magnesium slurry fuel are presented.
The ram jet accelerated from a Mach number of 1.75
to a Mach number of 3.48 in 15.5 seconds. During
this period a maximum acceleration of 4.6g was
obtained. Maximum values of air specific impulse
and gross thrust coefficient were calculated to be
151 seconds and 0.658, respectively. The rocket
gas generator used as a fuel-pumping system
operated successfully.


* AVAILABLE ON LOAN ONLY









DECLASSIFIED NACA RESEARCH REPORTS


NACA RM L55K09

WIND-TUNNEL INVESTIGATION AT TRANSONIC
SPEEDS OF A JET CONTROL ON A 350 SWEPT
WING. TRANSONIC-BUMP METHOD. Raymond D.
Vogler and Thomas R. Turner. February 1956.
17p. diagrs. (NACA RM L55K09)

Results are presented of a wind-tunnel investigation
at transonic speeds and at angles of attack of -40 to
160 to determine the characteristics of a jet control
on a 350 swept-semispan wing. The control con-
sisted of numerous holes normal to the wing surface
located on the 65-percent chord line and extending
in a spanwise direction from 0.133 to 0.70 of the
semispan. Force and moment data were obtained
by using ejected air with pressure ratios as high as
9.7 to 1 between the total pressure in the jets and
free-stream static pressure.


NACA RM L55K11

EXPERIMENTAL INVESTIGATION AT HIGH SUB-
SONIC SPEED OF THE ROLLING STABILITY
DERIVATIVES OF A COMPLETE MODEL HAVING
A CLIPPED-DELTA WING AND A HIGH HORIZON-
TAL TAIL. William C. Sleeman, Jr., and Albert G.
Few, Jr. February 1956. 32p. diagrs., tab.
(NACA RM L55K11)

Rolling-stability derivatives are presented for a
complete model having an aspect-ratio-3 clipped-
delta wing and a high horizontal tail for a Mach num-
ber range of 0.60 to 0.92 and for an angle-of-attack
range from 00 to a maximum of 130 for the lower
Mach numbers. The wing and delta plan-form hori-
zontal tail were swept back 450 at the leading edge
and had NACA 65A006 airfoil sections. Breakdown
tests were made to determine contributions of the
various component parts to the rolling-stability
derivatives.


NACA RESEARCH ABSTRACTS
AND RECLASSIFICATION NOTICE NO. 122


NACA RM L56C06

PREFLIGHT AND FLIGHT-TEST INVESTIGATION
OF A 50-PERCENT-MAGNESIUM 50-PERCENT
JP-4 SLURRY FUEL IN A TWIN-ENGINE RAM-JET
VEHICLE. Otto F. Trout, Jr., and Thomas L.
Kennedy. May 1956. 27p. diagrs., photos.
(NACA RM L56C06)

Performance data obtained from the preflight and
flight tests of a 50-percent-magnesium 50-percent
JP-4 slurry fuel are presented. Using this fuel,
the ram-jet vehicle accelerated from a Mach number
of 2.23 to 2.56 while climbing from an altitude of
3,320 to 37,940 feet. Data are presented for the
values of air specific impulse and gross thrust
coefficients over the range of altitudes tested.









NACA RM L56I24a

FLIGHT INVESTIGATION OF A RAM JET BURNING
MAGNESIUM SLURRY FUEL AND HAVING A CONI-
CAL SHOCK INLET DESIGNED FOR A MACH NUM-
BER OF 4.1. Walter A. Bartlett, Jr., and
Charles F. Merlet. January 1957. 23p. diagrs.,
photos. (NACA RM L56I24a)

A ram-jet vehicle having a conical shock inlet diffu-
ser designed for a Mach number of 4.1 and utilizing
a magnesium slurry fuel was tested in free flight.
Data are presented which show that the vehicle ac-
celerated from a Mach number of 1.73 to a maximum
Mach number of 3.84 in 13.2 seconds. During this
period a maximum acceleration of 6.1g was meas-
ured. Maximum values of air specific impulse of
150 seconds and gross thrust coefficient of 0.76 were
calculated.


* AVAILABLE ON LOAN ONLY


NACA Langley Field. Va.




NATIONAL ADVISORY COMMITTEE for AERONAUTICS
Division of Research Information
1512 H St., N. W. Washington 25, D. C.

REQUEST FORM RESEARCH PUBLICATIONS

1. Iftitles ofclassified publications are given, classify this form.
2. Please give complete NACA code number: for example. TN 1000, TM 1000. RhU L55B34a, N-1000.
(Non-NACA items listed in Research Abstracts should be ordered by N-number). Use of non-NACA
numbers may delay your request.
3. Current NACA Reports (formerly called Technical Reports) are available from the Superintendent of
S Documents, Goverwnment Printing Office. Washington 25. D. C.
4. Items not available for retention will be sent on loan if requested.

Item NACA First Author Title or other identification.
Code Number (Provide trh if you do not bare the complete NACA code number/

1

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3

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7

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9

10

Check one: -- Send loan items. Do not send loan items.
For NACA Use
Items not supplied Reason
Not available for circulation
Security classified
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Available from Superintendent of Documents,
Government Printing Office, Washington 25,
D. C.

--------------------------------- -------------------- ----------------------------
For NACA Use
Date

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Organization

Street Address

City, Zone, State
NACA Form 389 (March 1957)








UNIVERSITY OF FLORIDA


3 1262 08155 183 9




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