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

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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:
AA00009235:00052

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ional Advisory Committee for Aeronautics


S.


Research Abstracts


-.--CeRRENT NACA REPORTS

NACA Rept. 1176

DETERMINATION OF MEAN CAMBER SURFACES
FOR WINGS HAVING UNIFORM CHORDWISE LOAD-
ING AND ARBITRARY SPANWISE LOADING IN
SUBSONIC FLOW. S. Katzoff, M. Frances Faison
and Hugh C. DuBose. 1954. ii, 17p. diagrs., tab.
(NACA Rept. 1176. Formerly TN 2908)

Methods involving integration around the wing
boundary are presented for computing mean camber
surfaces to support uniform chordwise loading, with
either uniform or nonuniform spanwise loading, in
subsonic flow. For polygonal wings with uniform
area loading, analytical expressions are developed
for both the slopes and the ordinates of the mean
camber surfaces. Mean camber surfaces of several
calculated wings are shown.


NACA Rept. 1178

CALIBRATION OF STRAIN-GAGE INSTALLATIONS
IN AIRCRAFT STRUCTURES FOR THE MEASURE-
MENT OF FLIGHT LOADS. T. H. Skopinski,
William S. Aiken, Jr. and Wilber B. Huston. 1954.
ii, 29p. diagrs., 10 tabs. (NACA Rept. 1178.
Formerly TN 2993; RM L52G31)

A general procedure is developed for calibrating
strain-gage installations in aircraft structures for
application to flight measurements of loads. The
basic procedure can be modified as necessary to
suit the requirements of any particular structure.
The application of the procedure is illustrated by
results for two typical structures.


I. NAA Rept. 1183

SUPERSONIC FLOW PAST OSCILLATING AIRFOILS
INCLUDING NONLINEAR THICKNESS EFFECTS.
Milton D. Van Dyke. 1954. ii, 17p. diagrs. (NACA
Rept. 1183. Formerly TN 2982)

A solution to second order in thickness is derived
for harmonically oscillating two-dimensional airfoils
in supersonic floJ. For slow oscillations of an
arbitrary profile, the result is found as a series in-
cluding the third power of frequency. For arbitrary
frequencies, the n.ethod of solution for any specific
profile is indicated, and the explicit solution derived
for a single wedge. Nonlinear thickness effects are
found generally to reduce the torsional damping, and
so enlarge the range of Mach numbers within which
torsional instability is possible.



*AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW.,
THE REPORT TITLE AND AUTHOR.

O29 /o&2B-
t &9l-


NACA Rept. 1184

THE NORMAL COMPONENT OF THE INDUCED
VELOCITY IN THE VICINITY OF A LIFTING ROTOR
AND SOME EXAMPLES OF ITS APPLICATION.
Walter Castles, Jr. and Jacob Henri De Leeuw,
Georgia Institute of Technology. 1954. 11, 15p.
diagrs., 3 tabs. (NACA Rept. 1184. Formerly
TN 2912)
A method is presented for computing the approximate
values of the normal component of the induced veloc-
ity at points in the flow field of a lifting rotor. Tables
and graphs of the relative magnitudes of the normal
component of the induced velocity are given for se-
lected points in the longitudinal plane of symmetry of
the rotor and on the lateral rotor axis. A method is
also presented for using the tables and graphs to
determine the interference induced velocities arising
from the second rotor of a tandem- or side-by-side-
rotor helicopter and the induced flow angle at a
horizontal tail plane.

NACA Rept. 1188

ON THE USE OF THE INDICIAL FUNCTION CON-
CEPT IN THE ANALYSIS OF UNSTEADY MOTIONS
OF vINGS AND WING-TAIL COMBINATIONS.
Murray Tobak. 1954. iii, 43p. diagrs. (NACA
Rept. 1188)

The concept of indicial aerodynamic functions is
applied to the analysis of the short-period pitching
mode of aircraft. By the use of simple physical
relationships associated with the indicial-function
concept, qualitative studies are made of the separate
effects on the damping in pitch of changes in Mach
number, aspect ratio, plan-lorm shape, and fre-
quency. The concept is further shown to be of value
in depicting physically the induced effects on a tail
surface which follows in the wake of a starting for-
ward surface. Considerable effort is devoted to the
development of theoretical techniques whereby the
transient response in lift at the tail to the wing wake
may be estimated. Numerical results for several
representative cases are presented, and these are
analyzed to reassess the importance of the contribu-
tion to the rotary damping moment of the interfer-
ence lift at the tail.

NACA Rept. 1194

A STUDY OF HYPERSONIC SMALL-DISTURBANCE
THEORY. Milton D. Van Dyke. 1954. ii, 21p.
diagrs. (NACA Rept. 1194. Formerly TN 3173)

The small-disturbance equations are derived for
inviscid flow past thin bodies at high supersonic
speeds. Reinterpreted, they apply throughout the




WASHINGTON 25, D C.. CITING CODE NUMBER ABOVE EACH TITLE.






2




supersonic range. The theory is used to find pres-
sures on cones and wedges, initial gradients on
ogives, and initial pressure curvatures on ogives of
revolution. Additional approximations from existing
theories are discussed.




NACA Rept. 1196

AN ANALYTICAL STUDY OF THE EFFECT OF
AIRPLANE WAKE ON THE LATERAL DISPERSION
OF AERIAL SPRAYS. Wilmer H. Reed, m. 1954.
ii, 16p. diagrs., 3 tabs. (NACA Rept. 1196.
Formerly TN 3032)

An analysis is made to determine the trajectories
and deposit of aerial spray droplets which are issued
into the air disturbances generated by an agricul-
tural airplane. Various nozzle arrangements and
droplet-size spectra are considered with a view to
improving the uniformity and effective width of the
deposit.





NACA RM 55F08

METAL-BONDING ADHESIVES FOR HIGH-
TEMPERATURE SERVICE. John M. Black and R. F.
Blomquist, Forest Products Laboratory. July 1955.
22p. tab. (NACA RM 55F08)

Results of an investigation made for the purpose of
developing a metal-bonding adhesive with improved
heat-resistant properties are reported. The most
promising results were obtained with a formulation
of a phenol resin and an epoxy resin with certain heat
stabilizers and catalysts. An improved straight
epoxy-resin adhesive that has superior strength
properties at 2500 to 3000 F compared with any other
known epoxy-resin adhesive formulation is also
reported.





NACA TN 3298

A LOW-DENSITY WIND-TUNNEL STUDY OF
SHOCK-WAVE STRUCTURE AND RELAXATION
PHENOMENA IN GASES. F. S. Sherman, University
of California. July 1955. 83p. diagrs., photos.,
2 tabs. (NACA TN 3298)

The profiles and thicknesses of normal shock waves
of moderate strength were determined in terms of
the variation of the equilibrium temperature of an
insulated transverse cylinder in free-molecule flow.
The shock waves were produced in a steady state in
the jet of a low-density wind tunnel at initial Mach
numbers of 1.72 and 1.82 in helium and 1.78, 1.85,
1.90, 1.98, 3.70, and 3.91 in air. The shock thick-
ness was determined from the maximum slope of the
cylinder temperature profile. The experimental
shock profiles are compared with various theoreti-
cal predictions.


NACA
RESEARCH ABSTRACTS NO. 88



NACA TN 3470

GUST-TUNNEL INVESTIGATION OF THE EFFECT
OF A SHARP-EDGE GUST ON THE FLAPWISE
BLADE BENDING MOMENTS OF A MODEL HELI-
COPTER ROTOR. Domenic J. Maglieri and Thomas
D. Reisert.August 1955. 24p. diagrs., photos.
(NACA TN 3470)

Preliminary investigations have been made in the
Langley gust tunnel to determine the effects of a
sharp-edge vertical gust on the blade flapwise vibra-
tory bending moments of small model rotors having
either fixed-at-root or teetering lDades. Both rotor
configurations were tested up to a tip-speed ratio of
about 0.35. The results for simulated forward flight
indicate that the effect of the gust on the maximum
vibratory bending moments is of less importance for
the teetering rotor than for the fixed-at-root rotor. *
Increasing the rotor speed decreases the magnitude
of the vibratory bending moments resulting from a
given gust. At a given rotor speed, the magnitude of
the vibratory components due to the gust increases
with increasing tip-speed ratio. Increasing the rotor
speed at a constant forward velocity decreases the
maximum vibratory bending moments for all condi-
tions tested.





NACA TN 3471

THEORETICAL ANALYSES TO DETERMINE
UNBALANCED TRAILING-EDGE CONTROLS
HAVING MINIMUM HINGE MOMENTS DUE TO
DEFLECTION AT SUPERSONIC SPEEDS. Kennith L.
Goin. August 1955.52p. diagrs., tab. (NACA TN 3471.
Formerly RM L51F19)

Theoretical analyses have been made to determine
the plan forms of unbalanced trailing-edge flap-type
controls having minimum hinge moments due to
deflection and requiring minimum work to overcome
the hinge moments due to deflection at supersonic
speeds. Ratios of lift and rolling moment to hinge
moment and lift 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 number
have been considered.




NACA TN 3473

EFFECTS OF SWEEP AND ANGLE OF ATTACK ON
BOUNDARY-LAYER TRANSITION ON WINGS AT
MACH NUMBER 4.04. Robert W. Dunning and
Edward F. Ulmann. August 1955. 31p. diagrs.,
photos. (NACA TN 3473)

Wind-tunnel tests were conducted at Mach number
4.04 to determine the effects of leading-edge sweep,
angle of attack, and leading-edge thickness on the
boundary-layer transition on flat-plate wings. In
addition, some results were obtained on wings hav-
ing rounded leading-edge sections. Transition
points were determined for angles of attack up to






NACA
RESEARCH ABSTRACTS NO. 88



200 and for leading-edge sweep angles from 00
to 720. A correlation has been obtained, for the
present data, of the effects of leading-edge sweep
and angle of attack on boundary-layer transition on
plane surfaces when the equivalent stagnation pres-
sure (stagnation pressure calculated from the local
static pressure and the local Mach number normal
to the leading-edge) and the rate of change of the
transition Reynolds number with equivalent stagna-
tion pressure are considered.




NACA TN 3482

SUPPLEMENTARY CHARTS FOR ESTIMATING
PERFORMANCE OF HIGH-PERFORMANCE HELI-
COPTERS. Robert J. Tapscott and Alfred Gessow.
July 1955. 31p. diagrs. (NACA TN 3482)

Charts published in NACA TN 3323 for estimating
the performance of high-performance helicopters
were applicable to rotors having hinged rectangular
blades *ith a linear twist of -80. Supplementary
charts are presented herein covering twists of 00
and -160.




NACA TN 3484

ON SPECTRAL ANALYSIS OF RUNWAY ROUGHNESS
AND LOADS DEVELOPED DURING TAXIING. John
C. Houbolt, James H. Walls and Robert F. Smiley.
July 1955. 9p. diagrs. (NACA TN 3484)

The application of the technique of generalized har-
monic analysis as a means for determining airplane
taxiing loads is considered in a cursory manner.
Some results on runway roughness are reviewed and
some results obtained from taxiing tests of a large
airplane are given. An elementary extrapolation of
results for low taxiing velocities to higher velocities
is shown to be conservative. Also, oleo-strut fric-
tion is shown to be a very important factor. With
regard to the load-prediction phase of taxiing loads
by spectral techniques, much additional work is
required, especially with respect to the treatment
of the transfer function.




NACA TN 3487

ACOUSTIC RADIATION FROM TWO-DIMENSIONAL
RECTANGULAR CUTOUTS IN AERODYNAMIC
SURFACES. K. Krishnamurty, California Institute
of Technology. August 1955. 33p. diagrs., photos.
(NACA TN 3487)

The acoustic radiation from two-dimensional rec-
tangular cavities cut into a flat surface was investi-
gated by schlieren, hot-wire, and optical interfer-
ometric techniques. The relation of the frequency,
as measured in the gap by a hot-wire, to the gap
breadth at a particular Mach number and free-
stream temperature is shown. The intensity of the
radiation was measured by means of an optical
interferometer.


3



NACA TN 3488

SOME MEASUREMENTS OF FLOW IN A RECTAN-
GULAR CUTOUT. Anatol Roshko, California
Institute of Technology. August 1955. 21p. diagrs.
(NACA TN 3488)

The flow in a rectangular cavity, or slot, in the floor
of a wind tunnel is described by the results of pres-
sure and velocity measurements. Pressure distri-
butions on the cavity walls as well as measurements
of friction are presented. The effects of varying
depth-breadth ratio are shown.





NACA TN 3507

PRACTICAL CONSIDERATIONS IN SPECIFIC
APPLICATIONS OF GAS-FLOW INTERFEROMETRY.
Walton L. Howes and Donald R. Buchele. July 1955.
ii, 95p. diagrs., photos. (NACA TN 3507)

By extending the analysis in NACA TN 3340, equa-
tions which simultaneously account for refraction and
corner effects are derived for evaluating one-
dimensional density fields from optical interfero-
grams. The random error in measuring fringe
shifts is determined. Spurious surface phenomena
are discussed and a simple method for model aline-
ment is described. A method for determining the
density at a surface is described. Systematic and
random errors are computed for three representa-
tive experimental situations, namely, boundary
layers associated with supersonic and subsonic flows
along flat plates and free convection of heat from a
horizontal cylinder. Reasons for discrepancies
between theory and experiment are offered.



NACA TN 3515

ANALYSIS OF TWO-DIMENSIONAL COMPRESSIBLE-
FLOW LOSS CHARACTERISTICS DOWNSTREAM OF
TURBOMACHINE BLADE ROWS IN TERMS OF
BASIC BOUNDARY-LAYER CHARACTERISTICS.
Warner L. Stewart. July 1955. 48p. diagrs. (NACA
TN 3515)

Equations are derived for obtaining compressible-
flow boundary-layer characteristics for a simple
power velocity profile. Loss coefficients at the
trailing edge and after mixing are then obtained in
terms of these boundary-layer characteristics. This
analysis is used to study the general effect of com-
pressibility on blade-exit loss characteristics, the
significance of mixing downstream of the blade row,
and the effect of trailing-edge thickness on overall
loss coefficients.



NACA TN 3516

SUMMARY EVALUATION OF TOOTHED-NOZZLE
ATTACHMENTS AS A JET-NOISE-SUPPRESSION
DEVICE. Warren J. North. July 1955. 19p.
diagrs., photo. (NACA TN 3516)

Toothed attachments to a full-scale turbojet nozzle






4



were investigated for possible jet-noise reduction
and thrust penalty. The attachments caused slight
reductions in total sound power that are insignificant
when evaluated in terms of engine thrust losses and
aircraft payload penalty. At the reduced thrust
levels obtained with the toothed nozzles, correspond-
ing sound power reductions could be realized by
throttling the standard turbojet engine. Sound pres-
sure levels were reduced behind the engine but were
increased elsewhere. A typical resident in the air-
port neighborhood would hear increased loudness in
the middle frequencies; however, the resulting over-
all auditory effect is considered to be negligible.




NACA TN 3518

ROTATING-STALL CHARACTERISTICS OF A
ROTOR WITH HIGH HUB-TIP RADIUS RATIO.
Eleanor L. Costilow and Merle C. Huppert. August
1955. 59p. diagrs., photos. (NACA TN 3518)

The rotating-stall characteristics of a 0.9 hub-tip
ratio rotor are investigated. The test setup ap-
proximates the two-dimensional model upon which
recent theories have been based. Measurements of
pressure, temperature, and flow fluctuations during
stalled operation are reported. Experimental stall-
propagation rates are compared with those predicted
by theory.





NACA TN 3520

FLAME PROPAGATION LIMITS OF PROPANE AND
n-PENTANE IN OXIDES OF NITROGEN. Riley O.
Miller. August 1955. 29p. diagrs., 3 tabs. (NACA
TN 3520)

Flame propagation limits with propane and n-pentane
in oxides of nitrogen were obtained at subatmospheric
pressures in a 2-inch-diameter, 48-inch-length tube.
Three oxidants were investigated; namely, nitric
oxide NO, nitrogen tetroxide N204, and a nearly
equimolar mixture of these two oxides. Flames
propagate through all these mixtures with the com-
position limits occurring at equivalence ratios of
roughly 1/3 and 3. The minimum propagation pres-
sure of the fuel-NO mixtures in the 2-inch-diameter
tube were appreciably greater than that of the fuel-
N204 mixtures. In general, the data attest to the
relative chemical stability of NO in the hydrocarbon
flames.




NACA TN 3521

A COMPARISON OF THE MEASURED AND PRE-
DICTED LATERAL OSCILLATORY CHARACTERIS-
TICS OF A 350 S.VEPT-WING FIGHTER AIRPLANE.
Walter E. McNeill and George E. Cooper. August
1955. 22p. diagrs., photo., 3 tabs. (NACA
TN 3521. Formerly RM A51C28)


NACA
RESEARCH ABSTRACTS NO.88


Flight measurements of the lateral oscillatory char-
acteristics 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-
sideslip amplitude, 1

tions of Mach number for each test altitude, and are
compared with values computed from wind-tunnel
data and estimated stability derivatives.





NACA TN 3540

A REEVALUATION OF GUST-LOAD STATISTICS
FOR APPLICATIONS IN SPECTRAL CALCULA-
TIONS. Harry Press and May T. Meadows.
August 1955. 19p. diagrs. (NACA TN 3540)

The available information on the spectrum of atmos-
pheric turbulence is briefly reviewed. On the basis
of these results, a method is developed for the con-
version of available gust statistics normally given in
terms of counts of gust peaks into a form appropriate
for use in spectral calculation. The fundamental
quantity for this purpose appears to be the probabil-
ity distribution of the.rool-mean-square gust veloc-
ity. Estimates of the variation of this distribution
with altitude and weather condition are also derived
from available gust statistics.







DECLASSIFIED NACA REPORTS








NACA RM A51G13

AERODYNAMIC CHARACTERISTICS OF THE NACA
RM-10 RESEARCH MISSILE IN THE AMES I- BY
3-FOOT SUPERSONIC WIND TUNNEL NO. 2 -
PRESSURE AND FORCE MEASUREMENTS AT
MACH NUMBERS OF 1.52 AND 1.98. Edward W.
Perkins, Forrest E. Gowen, and Leland .H.
Jorgensen. September 1951. 37p. diagrs. (NACA
RM A51G13) (Declassified from Confidential,
7/20/55)
The pressure distribution along the body of zero
angle of attack and around the body for various lon-
gitudinal stations and for angles of attack to 150 are
presented. These data are compared with the method
of characteristics and with linear and second-order
theories. Aerodynamic coeflicients for the body
alone and the body-tail combination were obtained at
Reynolds numbers, based on model length, of 8.6
and 17.4 millions. The body-alone aerodynamic co-
efficients are compared with the theory of NACA
RM A9126 and with potential theory.






NACA
RESEARCH ABSTRACTS NO.88


NACA RM A55C07

FLIGHT TESTS OF LEADING-EDGE AREA SUC-
TION ON A FIGHTER-TYPE AIRPLANE WITH A
350 SWEPTBACK WING. Richard S. Bray and
Robert C. Innis. June 1955. 30p. diagrs.. photos.,
tab. (NACA RM A55C07) (DeclassLfied from
Confidential, 7 20. 55)

Flight measurements made of the low-speed lift
characteristics of an F-86 airplane equipped with a
porous leading-edge installation confirmed the re-
sults obtained from mind-tunnel tests of a similar
installation. Large gains in maximum lift coef-
ficient were obtained. Stalling characteristics were
tolerable to good. The porous leading edge had no
marked effect on the operational performance of the
airplane.






NACA RM E51F28

SOME EFFECTS OF BLADE TRALLING-EDGE
THICKNESS ON PERFORMANCE OF A SINGLE-
STAGE AXIAL-FLOW COMPRESSOR. J. J. Moses
and G. K. Serovy. October 1951. 14p. diagrs.,
tab. (NACA RM E51F28) (Declassified from
Confidential, 7 20, 55)

A set of modified NACA 65-series blower blades
designed for axial inlet velocity, high inlet Mach
number, and high blade loading was investigated for
trailing-edge thicknesses of 0.015, 0.030, and 0.045
inch to determine the effect of trailing-edge thick-
ness on single-stage axial-flow-compressor per-
formance. Trailing-edge thickness effects were
small except at the highest tip speed investigated
(915 ft/sec). Trailing-edge thicknesses up to 30
percent of maximum blade thickness were used with-
out sacrifice of performance of NACA 65-series
blades.






NACA RM E52K13

PRELIMINARY INVESTIGATION IN J33 TURBOJET
ENGINE OF SEVERAL ROOT DESIGNS FOR
CERAMAL TURBINE BLADES. George C. Deutsch,
Andre J. Meyer, Jr., and William C. Morgan.
January 1953. 24p. diagrs., photos.. 2 labs.
(NACA RM E52K13) (Declassified from Confidential,
7,20/55)
The practicability of using ceramals with compara-
tively low strategic material content for the blades
of aircraft turbines was determined in an experi-
mental investigation. Four blade root configurations
were examined. The most favorable results were
obtained for ceramal turbine blades with single ser-
ration interlock and dovetail root configurations.
Six of the interlock type were operated in a J33-A-33
engine for 68 hours, 23 minutes at rated service
speed and six of the dovetail type were operated for
58 hours, 28 minutes at rated speed. This result
confirms the conclusions of previous static design
studies.


5



NACA RM E53B27

AN INVESTIGATION OF HIGH-FREQUENCY COM-
BUSTION OSCILLATIONS IN LIQUID-PROPELLANT
ROCKET ENGINES. Adelbert O. Tischler,
Rudolph V. Massa, and Raymond L. Manlier. June
1953. 37p. diagrs., photos. (NACA RM E53B27)
(Declassified from Confidential, 7, 20, 55)

An experimental investigation of high-frequency
combustion oscillations (screaming) was conducted
with a 100-pound-thrust acid-hydrocarbon rocket
engine and a 500-pound-thrust oxygen-fuel rocket
engine. The oscillation frequencies could be cor-
related as a linear function of the parameter C', L,
where C* is the experimentally measured char-
acteristic velocity and L is the combustion-
chamber length. The tendency of the engines to
scream increased as chamber length was increased.
With engine configurations that normally had a low
efficiency, screaming resulted in increased per-
formance; at the same time, a five to tenfold in-
crease in heat-transfer rate occurred. It was pos-
sible, however, to achieve good performance without
screaming.





NACA RM E53G02

PRELIMINARY INVESTIGATION OF SEVERAL
ROOT DESIGNS FOR CERMET TURBINE BLADES
IN TURBOJET ENGINE. I ROOT DESIGN
ALTERATIONS. A. J. Meyer, Jr., G. C. Deutsch,
and W. C. Morgan. October 1953. 34p. photos.,
diagrs., tab. (NACA RM E53G02) (Declassified
from Confidential, 7,20'55)

Twelve engine evaluation tests of cermet turbine
blades are reported herein. The first five runswere
similar to the 10 engine tests reported previously.
The cermet blades were then redesigned on the basis
of information gained during the earlier runs. The
profile of the root was changed; the airfoil was re-
located on the platform and root; the root center
line was rotated with respect to the turbine axis;
and an attempt was made to improve the blade
material. These alterations primarily accomplished
a reduction in root bending stresses and resulted in
greatly increased blade life. The modified blades
were operated under increased power output from
the engine and were subjected to cyclic conditions
between take-off speed and idling speed.




NACA RM E53L15b

PRELIMINARY INVESTIGATION OF SEVERAL
TARGET-TYPE THRUST-REVERSAL DEVICES.
Fred W. Steffen, H. George Krull, and Carl C.
Ciepluch. March 1954. 43p. diagrs., tab. (NACA
RM E53L15b) (Declassified from Confidential,
7,20/55)

Thrust-reversal performance of several basic
target-type jet deflectors of various size and with
various modifications was obtained with unheated air
over a range of exhaust-nozzle pressure ratios from
1.7 to 3.0. Reversed thrusts up to 58 percent were
achieved. Maximum, or near maximum, reversal





6


for any deflector occurred at a spacing which was
mechanically feasible and which did not affect noz-
zle air flow. Two easily retractable deflectors, a
simple hemisphere and a double-skin hemisphere,
the latter designed to control reversed jet direction,
were investigated while installed on a model of a
turbojet engine tail pipe and cowling assembly.
Reversal produced by the installation with either
deflector was found to be about 40 percent at a noz-
zle pressure ratio of 2.0. With the simple hemi-
sphere, 50 percent of the available thrust was lost
through insufficient turning and 11 percent was lost
through air-flow reduction, shock, and friction.


NACA RM L7C25

AERODYNAMIC MEASUREMENTS MADE DURING
NAVY INVESTIGATION OF HUMAN TOLERANCE
TO WIND BLASTS. Donald L. Loving. March 11,
1947. 34p. diagrs., photos., 2 tabs. (NACA
RM L7C25) (Declassified from Confidential,
6/10/55)

Tests were made to obtain direct evidence of the
forces involved when the human head is suddenly
thrust into a rapidly moving air stream, as is the
case in bail-outs from aircraft at high speeds, and
also to determine the maximum speed considered
safe for the unprotected face to be exposed to wind
blasts. Included are the aerodynamic results of the
pressure measurements over a dummy forehead and
the air-blast forces on the unprotected heads tested
as obtained through the use of a strain gage attached
to the head rest of the test apparatus.



NACA RM L8G21

MEASUREMENTS OF THE CHORDWISE PRESSURE
DISTRIBUTIONS OVER THE WING OF THE XS-1
RESEARCH AIRPLANE IN FLIGHT. De E. Beeler,
Milton D. McLaughlin, and Dorothy C. Clift.
August 4, 1948. 35p. diagrs., photo., tab. (NACA
RM L8G21) (Declassified from Confidential,
7/20/55)

Presents results of chordwise pressure-distribution
measurements at a section near the midspan of the
left wing of the XS-1 research airplane with an 8-
percent-thick wing for a Mach number range of 0.75
to 1.25 at a normal-force coefficient of about 0.33
and for normal-force coefficients up to 0.93 at a
Mach number of about 1.16.




NACA RM L51F27

WATER LANDING INVESTIGATION OF A HYDRO-
SKI MODEL AT BEAM LOADINGS OF 18.9 AND 4.4.
Sidney A. Batterson. September 1951. 54p. diagrs.,
photos., tab. (NACA RM L51F27) (Declassified
from Confidential, 7/20/55)


NACA
RESEARCH ABSTRACTS NO.88


Presents results obtained from an investigation
conducted in the Langley impact basin of a model
having a flat rectangular plamnng surface, a pulled-
up bow, and a simulated landing wheel. Runs were
made at beam loading of 18.9 and 4.4. The results
include plots showing the variation of the non-
dimensional loads and motions with both the wetted
length and flight-path angle.






NACA RM L54H09


DRAG MEASUREMENTS ON A 1, 6-SCALE, FINLESS,
STING-MOUNTED NACA RM -10 MISSILE IN FLIGHT
AT MACH NUMBERS FROM 1.1 TO 4.04 SHOWING
SOME REYNOLDS NUMBER AND HEATING EF-
FECTS. Robert 0. Piland. October 1954. 20p.
diagrs., photos. (NACA RM L54H09) (Declassified
from Confidential, 7/20/55)

A 1/6-scale, finless NACA RM-10 missile, sting-
mounted on a parent body, has been flight tested to
a peak Mach number of 4.04. Measurements of total
drag, base drag, and wall temperature were obtain-
ed. Reynolds numbers of 17 x 106 to 47 x 106, based
on body length, corresponding to Mach numbers 1.07
to 4.04 were encountered. Total- and base-drag
coefficients are correlated with wind-tunnel results.
The base-pressure coefficients are also compared
with calculations using Love's methods for the
estimation of base pressures. The friction drag is
estimated from test measurements and a calculated
pressure drag and is compared with theory.





NACA RM L54103

AN INVESTIGATION OFTHE CHARACTERISTICS
OF THE NACA RM-10 (WITH AND WITHOUT FINS)
IN THE LANGLEY 11-INCH HYPERSONIC TUNNEL
AT A MACH NUMBER OF 6.9. William D. McCauley
and William V. Feller. November 1954. 37p.
diagrs., photos. (NACA RM L54103) (Declassified
from Confidential, 7/20/55)

Pressure distributions on the body and force balance
results for the NACA RM-10 with and without fins
are presented for angles of attack from 0 to about
200 at M = 6.9. The experimental results are
compared with the following theories: method of
characteristics, Van Dyke's second-order theory,
conical-shock two-dimensional expansion theory,
Newtonian impact theory, Grimminger, Williams,
and Young's correlation prediction, and linear
theory. Interference between fins and body is dis-
cussed. The effect of Mach number on the compo-
nents of zero-lift drag and of the pitching moment,
center of pressure, and lift-curve slope at zero lift
are presented.






NACA LangleylHd'
I`?..





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