Research abstracts

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

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Subjects / Keywords:
Aeronautics -- Abstracts -- Periodicals   ( lcsh )
Aeronautics -- Research -- Abstracts -- Periodicals   ( lcsh )
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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:
AA00009235:00017

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National Advisory Committee For Aeronautics


Research Abstracts
NO.56


JANUARY 13, 1954


CURRENT NACA REPORTS

NACA Rept. 1099

AIR FORCES AND MOMENTS ON TRIANGULAR
AND RELATED WING WITH SUBSONIC LEADING
EDGES OSCILLATING IN SUPERSONIC POTENTIAL
FLOW. Charles E. Watkins and Julian H. Berman.
1952. il, 25p. diagrs., tab. (NACA Rept. 1099.
SFormerly TN 2457)

This analysis treats the air forces and moments in
supersonic potential flow on oscillating triangular
wings and a series of sweptback and arrow wings
with subsonic leading edges and supersonic trailing
edges. The linearized velocity potential for the
wings undergoing sinusoidal torsional oscillations
simultaneously with vertical translations is derived
to the form of a power series in terms of a frequency
parameter. Although as many terms of such a
series expansion as may be desired can be deter-
mined, the terms after the first few become very
cumbersome. Closed expressions that include the
reduced frequency to the third power, an order
which is sufficient for a large class of practical ap-
plications, are given for the velocity potential and
components of chordwise section force and moment
coefficients.


-.4
.i4ACA Rept. 1103

GENERALIZED THEORY FOR SEAPLANE IMPACT.
Benjamin Milwltzky. 1952. iii, 75p. diagre.,
4 tabs. (NACA Rept. 1103)

The motions, hydrodynamic loads, and pitching
moments experienced in impacts of V-bottom sea-
planes are analyzed and compared with experiment.
The analysis is presented in terms of generalized
variables which are related through a single param-
eter, called the approach parameter x. For use in
the design of seaplanes, charts are presented show-
ing the generalized relationships which apply
throughout the impact; charts are also presented
which show the variations with K of the generalized
variables at the instants of maximum acceleration,
maximum pitching moment about the step, maximum
penetration, and exit during rebound. The effects of
chine immersion on the maximum load are also de-
termined. Extensive experimental data are pre-
sented to permit evaluation of the theoretical re-,..
sults.


U'

(~ .jr*'


aJ


NACA Rept. 1118

METHOD FOR CALCULATION OF LAMINAR HEAT
TRANSFER IN AIR FLOW AROUND CYLINDERS OF
ARBITRARY CROSS SECTION (INCLUDING LARGE
TEMPERATURE DIFFERENCES AND TRANSPIRA-
TION COOLING). E. R. G. Eckert and John N. B.
Livingood. 1953. ii, 29p. diagrs. (NACA
Rept. 1118. Formerly TN 2733)

A method which permits approximation of local heat-
transfer coefficients in the laminar-flow region
around cylinders of arbitrary cross section from
those for wedge-type profiles is extended to Include
the effects of large temperature differences and
transpiration cooling. Charts prepared from exact
solutions of the laminar boundary-layer equations
for wedge-type profiles which allow for these effects
yield results with a minimum of calculation. Appli-
cation of the method to circular and elliptic tran-
spiration cooled cylinders is made to determine
local heat-transfer c pffitlt d surface tempera-
tures and to deters!fje Ya ti in coolant flow
required for m th a conit$ face tempera-
ture. '

NACA Rept. lI dDEC 9 195 .

RECIPROCITY I(ELATIONS IN AERODYNAMICS.
Max. A. Heasipt and John R. Sprelter.' 1953. ii,
16p. diagrs. (NACA'Re&LJ.U19. Formerly
TN 2700) -

Reserve-flow theorems in aerodynamics are shown
to be based on the same general concepts involved in
many reciprocity theorems in the physical sciences.
Reciprocal theorems for both steady and unsteady
motion are found as a logical consequence of this ap-
proach. No restrictions on wing plan form or flight
Mach number are made beyond those required in lin-
earized compressible-flow analysis. A number of
examples are listed, including general integral theo-
rems for lifting, rolling, and pitching wings and for
wings in nonuniform downwash fields. Correspond-
ence is also established between the build-up of cir-
culation with time of a wing starting impulsively from
rest and the build-up of lift of the same wing moving
In the reverse direction into a sharp-edge gust.


NACA TN 3030

A METHOD FOR CALCULATING THE SUBSONIC
STEADY-STATE LOADING ON AN AIRPLANE
WITH A WING OF ARBITRARY PLAN FORM AND
STIFFNESS. W. L. Gray and K. M. Schenk,
Boeing Airplane Company. December 1953. ii,
L 120p. diagrs., tab. (NACA TN 3030)


"AVAILABLE ON LOAN ONLY. ---
ADDRESS REQUESTS FOR DOCUMENT TONACA, 1724 F STNW; WASHINGTON 2s, D. C., CITING CODE NUMBER ABOVE EACH TITLE
THE REPORT TITLE AND AUTHOR. .-

6Zf./30/-5
2s-S I-








2

A method for computing the span load distribution on
an elastic airplane wing at specified airplane weights
and load factors is given. The method which applies
for suberitical Mach numbers for wings of arbitrary
plan form and stiffness distribution includes the ef-
fects of external stores and fuselage on the spanwise
loading. Modifications are outlined for treating
tail-boom and tailless airplane configurations and
for calculating the divergence dynamic pressure of a
swept wing with a large external store. A method
is also outlined for reducing wind-tunnel data to ob-
tain effective aerodynamic coefficients which are
Mach number can readily be evaluated from the
aerodynamic coefficients thus obtained.


NACA TN 3054

INVESTIGATION AT SUPERSONIC SPEEDS OF THE
WAVE DRAG OF SEVEN BOATTAIL BODIES OF
REVOLUTION DESIGNED FOR MINIMUM WAVE
DRAG. August F. Bromm, Jr. and Julia M.
Goodwin. December 1953 14p. diagrs., photo.
(NACA TN 3054)

Results are presented from an investigation of the
variation with Reynolds number and Mach number of
the wave drag of seven boattail bodies of revolution
designed for minimum wave drag according to the
theory presented in NACA TN 2550.


NACA TN 3057

A SIMPLIFIED MATHEMATICAL MODEL FOR
CALCULATING AERODYNAMIC LOADING AND
DOWNWASH FOR WING-FUSELAGE COMBINA-
TIONS WITH WINGS OF ARBITRARY PLAN FORM.
Martin Ziotruck and Samuel W. Robinson, Jr.
January 1954. 38p. diagrs. (NACA TN 3057.
Formerly RM L52J27a)

For the purpose of calculating the aerodynamic load-
ing on the fuselage, the midwing wing-fuselage com -
bination with a fuselage of circular cross section can
be represented by a simple system of horseshoe
vortices located on the wing with images located in-
side the fuselage. By using this simplified mathe-
matical model or the extension of it given in an ap-
pendix for nonmidwing configurations with fuselages
of arbitrary cross section, a method for calculating
the lift and longitudinal center of pressure on the
fuselage in the presence of the wing at subsonic
speeds is presented. In addition the report shows
how the simplified mathematical model can be used
for calculating the downwash behind the wing and for
calculating the spanwise lift distribution on the wing
for midwing configurations with axisymmetric fuse-
lages.


NACA TN 3060

USE OF ELECTRIC ANALOGS FOR CALCULATION
OF TEMPERATURE DISTRIBUTION OF COOLED
TURBINE BLADES. Herman H. Ellerbrock, Jr.,
Eugene F. Schum and Alfred J. Nachligall.
December 1953. 116p. diagrs., photos., 6 tabs.
(NACA TN 3060)


NACA
RESEARCH ABSTRACTS NO. 56

An investigation was conducted to develop simple,
inexpensive electric analogs for determining tem-
peratures of cooled turbine blades. The accuracy of
such analogs was determined by fabricating tree or
specific blade configurations and comparing values'eBf
blade temperatures obtained with them with calcu-
lated values of blade temperature when air-cooled
blades were considered and by relaxing the analog
values for liquid-cooled blades to determine whether
the residuals were acceptable. In general, good
agreement was achieved between calculated and
analog values of blade temperature.





NACA TN 3061

STRESSES AROUND RECTANGULAR CUTOUTS IN
TORSION BOXES. Paul Kuhn and James P.
Peterson. December 1953. 71p. diagrs., 2 tabs.
(NACA TN 3061)

A method is presented for calculating the stresses
produced by rectangular cutouts of any size in
torsion boxes. The problem is divided into a sbox
problem" and a "cover problem. The box problem
is a special case of the general method of analyzing
torsion boxes without cutouts. In the cover problem,
simple shear-lag theory is used to obtain "key'
stresses; the final distributions are obtained from
these key stresses by means of simple rules or
empirical distribution curves. Comparisons with
the results from three series of tests in which the
dimensions of the cutouts varied over a wide range
are shown.







NACA TN 3088

DETERMINATION OF THE FLYING QUALITIES OF
THE DOUGLAS DC-3 AIRPLANE. Arthur
Assadourian and John A. Harper. December 1953.
67p. diagrs., photos., tab. (NACA TN 3088)

Data are presented showing the longitudinal and
lateral stability and control characteristics and the
stalling behavior of the Douglas DC-3 airplane and
the compliance of these flying qualities with the cur-
rent Air Force-Navy specifications. Even though the
DC-3 was designed and built more than two decades
ago, the airplane satisfied most of the current speci-
fications for its type. However, the airplane had
static longitudinal instability for certain conditions of
airspeed and center-of-gravity position in the power-
on configurations, the specified maximum elevator
control-force gradient in maneuvers was exceeded in
most cases, the rudder forces required to overcome
the adverse aileron yaw were excessive and the
rudder and aileron forces in steady sideslips tended
to lighten at the higher sideslip angles. Typical
frequency-response characteristics of the airplane
are also presented for those interested in automatic
stabilization.








NACA
RESEARCH ABSTRACTS NO. 56

NACA TN 3091

FLOW PROPERTIES OF STRONG SHOCK WAVES
IN XENON GAS AS DETERMINED FOR THERMAL
EQUILIBRIUM CONDITIONS. Alexander P. Sabol.
December 1953. 29p. diagrs., photos., 2 labs.
(NACA TN 30911

The results of calculations are presented for the
purpose of showing the effects of ionization and elec-
tronic excitation on the flow properties of one dimen-
sional shock waves in xenon gas. The calculations
use statistical-mechanics theory with the assumption
that thermal equilibrium exists at all points in the
flow. The calculations are made by both including
.and neglecting the effects of electronic excitation,
and the results of the calculations are qualitatively
compared with available experimental data.



NACA TN 3092

HYDRODYNAMIC DRAG OF 12- AND 21-PERCENT-
THICK SURFACE-PIERCING STRUTS. ClaudeW.
Coffee, Jr. and Robert E. McKann. December 1953.
28p. diagrs., photos., tab. (NACA TN 3092)

An investigation has been made to determine the
hydrodynamic drag of three surface-piercing un-
tapered struts at approximately 0o angle of yaw at
depths up to 6 chords for speeds up to 80 fps at
various angles of rake. Two struts had NACA
661-012 airfoil sections, one with a 4-inch chord
and the other with an 8-inch chord. The third strut
had an NACA 664-021 airfoil section and a 4-inch
chord.




NACA TN 3094

THE RESISTANCE TO AIR FLOW OF POROUS
MATERIALS SUITABLE FOR BOUNDARY-LAYER-
CONTROL APPLICATIONS USING AREA SUCTION.
Robert E. Dannenberg, James A. Weiberg and
Bruno J. Gambucci. January 1954. 21p.
diagrs., photos., tab. (NACA TN 3094)

Measurements have been made of resistance to air
flow of commercially available porous materials.
Three general types of porous media were tested -
granular sinteredd metals), fibrous (felt cloths and
filter papers), and perforated. The flow-resistance
characteristics were ascertained over a wide range
of differential pressures across the material. The
flow resistance of one of the materials was deter-
mined for various values of air pressure at the up-
stream lace of the material.



NACA TN 3114

ANALYSIS OF MULTICELL DELTA WINGS ON
CAL-TECH ANALOG COMPUTER. Richard H.
MacNeal and Stanley U. Benscoter, California
Institute of Technology. December 1953. 84p.
diagrs., 6 tabs. (NACA TN 3114)


3


Using the Cal-Tech analog computer, structural
analyses have been made of two delta wings with 450
leading edges. One of these has a constant-depth
rectangular cross section and the other has a bicon-
vex cross section that is linearly tapered in the
spanwise direction. The wings extend through the
fuselage and are rigidly supported along two lines at
the faces of the fuselage. Deflections and all inter-
nal forces have been calculated for concentrated
static loads. Vibration modes are also presented.
The effects of neglecting shearing strains in the ribs
and spars and also of assuming the ribs to be rigid
have been investigated by modifying the electric cir-
cuits to correspond to these simplifications.




NACA RM E53129

EFFECTS OF ADDITIVES ON PRESSURE LIMITS OF
FLAME PROPAGATION OF PROPANE-AIR MIX-
TURES. Frank E. Belles and Dorothy M. Simon.
December 1953. 36p. diagrs., tab. (NACA
RM E53129)

Seven additives in 0.5-volume-percent concentration
were studied for their effects on the low -pressure
limits of flame propagation of propane-air mixtures.
The limits were measured in a flame tube of new de-
sign. Mixtures containing approximately 2 to 8 per-
cent propane by volume were studied. The limit
curves were without looes on the rich side and were
closely related to quenching-distance data measured
by th* flash-back of a Bunsen flame. The data were
analyzed by means of the experimental curves and
the Le Chatelier law governing the flammability
limits of mixed fuels. Ethyl nitrate and chloropicrin
were found to be definite promoters of flame propa-
gation in rich mixtures. Chloropicrin and methyl
bromide inhibited propagation in lean mixtures; it
was concluded that the effect is chemical and that
these additives do not act merely as inert gases.
None of the additives promoted flame propagation in
lean mixtures more than could be explained by the
contribution of the additive to the total fuel in the
mixture. Methyl bromide increased the minimum
pressure for flame propagation and was the only
additive that had an appreciable effect on the mini-
mum. Carbon disulfide had a large inhibitory effect
on flame propagation in lean mixtures, as defined by
deviations from the requirements of Le Chatelier's
law.

BRITISH REPORTS


N-27317*

Royal Aircraft Establishment (Gt. Brit.)
STRENGTH DATA ON CIRCULAR METAL CYLIN-
DERS UNDER AXIAL COMPRESSION. J. H.
Shelley. August 1953. 87p. diagrs., photos., 3
tabs. (RAE Tech. Note Structures 122)

Failure of circular metal cylinders under axial com-
pression is examined empirically for the practical
range of the wall thickness radius ratio. The
cylinders considered are those of length such that
end conditions do not have an appreciable effect upon








4

failure, but not long enough to fail as Euler struts.
The paper Is based on experimental results given In
15 separate publications, and covers wide ranges of
material specifications, manufacturing methods,
testing procedures, and specimen sizes. The de-
sign proposals submitted use as far as practicable
the formulas suggested by other authors, and are
compared with the experimental results obtained for
381 steel and 143 aluminum alloy specimens over the
full range of t/R.



MISCELLANEOUS

NACA Rept. 1096

Errata No. I on "EXPERIMENTAL DETERMINA-
TION OF THE EFFECT OF HORIZONTAL-TAIL
SIZE, TAIL LENGTH, AND VERTICAL LOCATION
ON LOW-SPEED STATIC LONGITUDINAL STA-
BILITY AND DAMPING IN PITCH OF A MODEL
HAVING 450 SWEPTBACK WING AND TAIL SUR-
FACES. Jacob H. Lichtenstein. 1952.



NACA Rept. 1099

Errata No. 1 on "AIR FORCES AND MOMENTS ON
TRIANGULAR AND RELATED WINGS WITH SUB-
SONIC LEADING EDGES OSCILLATING IN SUPER-
SONIC POTENTIAL FLOW. Charles E. Watkins
and Julian H. Berman. 1952.


NACA
RESEARCH ABSTRACTS NO. 56

NACA Rept. 1103

Errata No. 1 on GENERALIZED THEORY FOR
SEAPLANE IMPACT. Benjamin Milwitzkcy. 1952%..



N-27631

Advisory Group for Aeronautical Research and
Development. SPECIFICATIONS FOR AGARD
WIND TUNNEL CALIBRATION MODELS. (Pre-
sented at Rome AGARD conference,
December 16-17, 1952) 4p. diagram. (Advisory
Group for Aeronautical Research and Development.
AG4/M3)

Two standard wind-tunnel models accepted by the
AGARD are described. One is the NACA RM-10 on
which considerable data are available immediately.
The RM-10 is designated as AGARD calibration
model A. The other, AGARD calibration model B,
is a new configuration consisting of a body-wing
combination. The wing is a delta wing with a span
four times the body diameter. The wing consists of
an equilateral triangle with a symmetrical circular-
arc section. The body is a body of revolution.








NACA
RESEARCH ABSTRACTS NO. 56


DECLASSIFIED NACA REPORTS


THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM RESTRICTED TO
UNCLASSIFIED, 12 11,53.



NACA RM 51B05

PROPERTIES OF LOW-CARBON N-155 ALLOY BAR
STOCK FROM 12000 TO 18000 F. J. W. Freeman
and A. E. White, University of Michigan. May 3,
1951. 103p. diagrs., photos., 21 tabs. (NACA
RM 51B05)


Results are presented of an investigation undertaken
to establish rupture and total-deformation strengths
(Odesignr data) at 12000 to 18000 F for typical com-
mercial treatments of low-carbon N-155 alloy.
Rupture data are given for bar stock from two heats
at 12000, 13500, and 15000 F and for one of the heats
at 16500 and 18000 F in several conditions of prior
treatment. Because the variation between heats in-
dicated that the data would be of questionable value as
typical design data, the work was stopped with data on
stress and time for total deformation somewhat in-
complete and with very little duplication between the
two heats.



NACA RM 51D20

REVIEW OF CURRENT AND ANTICIPATED LUBRI-
CANT PROBLEMS IN TURBOJET ENGINES. NACA
Subcommittee on Lubrication and Wear. Appendix A:
THE DEVELOPMENT OF LUBRICANTS FOR EX-
TREME TEMPERATURE OPERATIONS FOR CUR-
RENT AND PROJECTED TURBINE-POWERED AIR-
CRAFT ENGINES. C. M. Murphy and W. A. Zisman.
Appendix B: AIRCRAFT TURBINE LUBRICANTS
FOR MILITARY USE. C. C. Singleterry. Appendix
C: BEARING PROBLEMS IN CURRENT AND PRO-
JECTED TURBO POWER PLANT. C. J. McDowall.
Appendix D: INVESTIGATIONS OF ROLLING BEAR-
INGS AT SKF INDUSTRIES, INC. H. Styri. Appen-
dix E: LUBRICANTS FOR OPERATION OF AIR-
CRAFT GAS TURBINE ENGINES OVER WIDE
TEMPERATURE RANGES. NACA Lewis Flight Pro-
pulsion Laboratory Staff. April 20, 1951. 46p.
diagram 4 tabs. (NACA RM 51D20)


Present and future lubricants for aircraft gas-
turbine engines using rolling contact bearings must
meet operating requirements of fludity, lubricity,
volatility, and stability over bearing operating tem-
perature regions of (1) -650 to 3500 F (current), (2)
-650 to 5000 F, and (3) -650 to 7500 F. The most
promising lubricants for these temperature ranges
involve the use of synthetics (that is, aliphatic
diesters) in whole or in part.


5


NACA RM 53B19

WEIGHT BAR CHARTS. B. J. Saelman and H. W.
Vick, Lockheed Aircraft Corporation. June 1953.
62p. diagrs., tab (NACA RM 53B19)


The weight breakdowns in bar chart lorm in the
present report have been prepared to give an
Indication of "how we spend our weight. "' They are
expected to focus attention on the design functions
that warrant weight reduction activity on future
models.





NACA RM A50A04a

CHORDWUB AND SPANWISE LWADINGS MEASURED
AT LAW PBEEDS ON A LARGE TRIANGULAR WING
HAVING AN ASPECT RATIO OF 2 AND A THIN,
SUBSONIC-TYPE AIRFOIL SECTION. David
Graham. March 13, 1950. 55p. diagrs., photo., 2
tabs. (NACA RM A50A04a)


Pressure data have been obtained on a large triangu-
lar wing having an aspect ratio of 2 and a modified
NACA 0005 section. The data were obtained at zero
angle of sideslip through an angle-of-attack range
from zero lift through the stall with three deflections
of a tralling-edge flap. Chordwise pressure distri-
butions, section lift characteristics, section centers
of pressure, and span load distribution are presented.
The tests were conducted at a dynamic pressure of
25 pounds per square foot resulting in a Reynolds
number, based on the mean aerodynamic chord, of
15.3 x 100. The Mach number was 0.13.





NACA RM A50A10

EFFECTS OF TWIST AND CAMBER ON THE LOW-
SPEED CHARACTERISTICS OF A LARGE-SCALE
450 SWEPT-BACK WING. Lynn W. Hunton.
March 20, 1950. 32P. diagrs., photos., tab. (NACA
RM A50A10)


Tests were made of a large-scale, semispan, swept-
back, wing-fuselage model cambered and twisted for
an ideal lift coefficient of 0.4 and of a similar model
but having no camber and no twist. The 0.25 chord
lines of the wings were swept back 450 and the wings
had an aspect ratio of 6.0 and a taper ratio of 0.5.
The aerodynamic characteristics, span loading, and
tuft-study results for the models are presented for a
Reynolds number of 8 million at a Mach number of
0.2. Predicted results from available theoretical
methods are compared with the experimental data.








6


NACA RM A50B13

PRESSURE COEFFICIENTS AT MACH NUMBERS
FROM 0.60 TO 0.85 FOR A SEMISPAN WING WITH
NACA 0012-64 SECTION, 20-PERCENT-CHORD
PLAIN AILERON, AND 0 AND 450 SWEEPBACK.
Walter J. Krumm. April 19, 1950. 36p. diagrs.,
30 tabs. (NACA RM A50B13iS)


The pressure measurements obtained during tests of
a semispan wing having the NACA 0012-64 section
are presented as coefficients. The measurements
were made at Mach numbers from 0.60 to 0.85 for
the wing with 00 and with 450 sweepback and with a
20-percent-chord, plain, trailing-edge aileron.




NACA RM A50E09

THE EFFECTS OF COMPRESSIBILITY ON THE
PRESSURES ON A BODY OF REVOLUTION AND ON
THE AERODYNAMIC CHARACTERISTICS OF A
WING-NACELLE COMBINATION CONSISTING OF
THE BODY OF REVOLUTION MOUNTED ON A
SWEPT-BACK WING. Frederick W. Boltz and
Benjamin H. Beam. July 26, 1950. 68p. diagrs..
photos., 2 tabs. (NACA RM A50E09)


Results of wind-tunnel tests are presented for a
semispan model of a wing-nacelle combination and
for an isolated body of revolution similar to the na-
celle. The wing had the leading edge swept back
37.250, an aspect ratio of 6.04, a taper ratio of 0.5,
zero twist, and NACA 641-212 sections. The simu-
lated nacelle and the body of revolution each had a
fineness ratio of 6.5. Lift, drag, pitching-moment,
and pressure data for the wing-nacelle combination
were obtained for Mach numbers from 0.18 to 0.92
at a Reynolds number of 2,000,000. Drag and pres-
sure data were obtained for the isolated body of
revolution for Mach numbers from 0.18 to 0.95 at a
Reynolds number of 3,700,000.




NACA RM A50H17

HIGH-SPEED AERODYNAMIC CHARACTERISTICS
OF A LATERAL-CONTROL MODEL. III SECTION
CHARACTERISTICS, FENCE STUDIES, AND TABU-
LATED PRESSURE COEFFICIENTS WITH MODI-
FIED NACA 0012-64 SECTION, 26.6-PERCENT-
CHORD, PLAIN AILERON, 00 AND 450 SWEEP-
BACK. Walter J. Krumm and Joseph W. Cleary.
November 22, 1950. 79p. diagrs., photos.. 4 tabs.
(NACA RM A50H17)


The spanwise variations of section normal-force and
pitching-moment coefficients were determined from
wind-tunnel pressure-distribution measurements of a
semispan wing with a modified NACA 0012-64 section
and a 26.6-percent-chord plain aileron. Tests were
made with the wing unswept and swept back 450 for
Mach numbers covering the force-divergence range
of the unswept wing. The effect of wing fences on


NACA
RESEARCH ABSTRACTS NO. 56


the static longitudinal stability of the swept wing was
investigated. Wing pressure coefficients arfre-
sented in tables. -




NACA RM A50H23

WIND-TUNNEL INVESTIGATION OF THE EFFECTS
OF A JET-ENGINE NACELLE ON THE AERO-
DYNAMIC CHARACTERISTICS OF A 37.250 SWEPT-
BACK WING AT HIGH SUBSONIC SPEEDS.
Frederick W. Boltz and Donald A. Buell.
October 24, 1950. 28p. diagrs., photos. (NACA
RM A50H23)


This report presents results of wind-tunnel tests of
a model of a jet-engine nacelle in combination with a
wing having the leading edge swept back 37.250 and
an aspect ratio of 6.04. The nacelle was mounted on
the lower surface of the wing with the air inlet
slightly behind the wing leading edge and normal to
the nacelle axis. Lift, drag, pitching moment, and
ram-recovery data are presented for Mach numbers
from 0.18 to 0.92 at a constant Reynolds number of
2,000,000. Surface pressure data are presented for
Mach numbers near that of drag divergence.




NACA RM A50J24

THE EFFECTS OF CAMBER AND TWIST ON THE
AERODYNAMIC LOADING AND STALLING CHARAC-
TERISTICS OF A LARGE-SCALE 450 SWEPT-BACK
WING. Lynn W. Hunton and Joseph K. Dew.
January 24, 1951. 40p. diagrs., photo., tab.
(NACA RM A50J24)


Pressure-distribution measurements were obtained
on a large-scale, semispan, sweptback, wing fuse-
Lage model cambered and twisted for a design lift
coefficient of 0.4 and on a similar model having no
camber and no twist. The wings had 450 sweepback
of the quarter-chord line, an aspect ratio of 6.0, and
a taper ratio of 0.5. Pressure distributions are
presented from tests made at a Reynolds number of
8 million and a Mach number of 0.2. Stalling char-
acteristics and spanwise distributions of local lift
coefficient and local center of pressure are derived
from the data. Theoretical span loadings are com-
pared with the experimental results.




NACA RM A50K28a

THE EFFECTS OF INCREASING THE LEADING-
EDGE RADIUS AND ADDING FORWARD CAMBER
ON THE AERODYNAMIC CHARACTERISTICS OF A
WING WITH 350 OF SWEEPBACK. Fred A. Demele
and Fred B. Sutton. February 9, 1951. 27p.
diagrs., photo., tab. (NACA RM A50K28a)








NACA
RESEARCH ABSTRACTS NO. 56

Presented in this report are results of an investiga-
tion of a sweptback wing which was modified by in-
creasing the leading-edge radius and introducing a
small amount of camber over the forward portion of
the chord. The wing had 350 of sweepback at the
quarter-chord line, and an aspect ratio of 4.5. The
original wing section was the NACA 64A010 normal
to the quarter-chord line. Aerodynamic character-
istics are presented for Reynolds numbers from
2,000,000 to 11,000,000 at a Mach number of 0.21 and
for Mach numbers from 0.21 to 0.94 at a Reynolds
number of 2,000,000. A comparison is made with
results of an investigation of the unmodified wing.


NACA RU A51C12a

INFLUENCE OF AIRFOIL TRAILING-EDGE ANGLE
AND TRAILING-EDGE-THICKNESS VARIATION ON
THE EFFECTIVENESS OF A PLAIN FLAP AT HIGH
SUBSONIC MACH NUMBERS. Albert D. Hemenover
and Donald J. Graham. May 22, 1951. 103p.
diagrs., photos., 5 tabs. (NACA RM A51C12a)


The effects of variation of trailing-edge angle and
trailing-edge thickness on the characteristics of a
10-percent-chord thick NACA airfoil section with a
25-percent-chord plain flap are appraised from
wind-tunnel tests at Mach numbers from 0.3 to 0.9
and Reynolds numbers varying correspondingly from
1 to 2 million. The airfoil trailing-edge angle was
varied from 180 to 60, and the trailing-edge thick-
ness from zero to the thickness at the flap hinge line.



NACA RM A51DO2


INVESTIGATION IN THE AMES 12-FOOT PRES-
SURE WIND TUNNEL OF A MODEL HORIZONTAL
TAIL OF ASPECT RATIO 3 AND TAPER RATIO 0.5
HAVING THE QUARTER-CHORD LINE SWEPT
BACK 450. Carl D. Kolbe and Angelo Bandettini.
June 25, 1951. 97p. diagrs., photo., 2 tabs.
(NACA RM A51D02)

The airfoil section of the model of a horizontal tail
was the NACA 64A010 in planes inclined 450 to the
plane of symmetry. The test Reynolds number was
varied from 2,000,000 to 18,000,000 at a Mach
number of 0.25 and the Mach number was varied from
0.25 to 0.94 at Reynolds numbers of 2,000,000 and
4,000,000. Lift, drag, pitching moment, hinge
moment, and the pressure difference across the
elevator-nose seal are reported.



NACA RM A51D18

EFFECTS OF DOUBLE-SLOTTED FLAPS AND
LEADING-EDGE MODIFICATIONS ON THE LOW-
SPEED CHARACTERISTICS OF A LARGF-SCALE
45o SWEPT-BACK WING WITH AND WITHOUT
CAMBER AND TWIST. Harry A. James and Joseph
K. Dew. July 1951. 37p. diagrs., photo., 3 tabs.
(NACA RM A51D18)


7

Tests were made on two large-scale semispan,
wing-fuselage models swept back 45 (aspect ratio
6.0, taper ratio 0.5) equipped with partial-span,
double-slotted flaps and various leading-edge modi-
fications. One wing was cambered and twisted for a
design lift coefficient of 0.4; the other wing had no
camber and no twist. The lift, drag, and pitching-
moment characteristics were determined at a
Reynolds number of 8 million. Theoretical predic-
tions of lift increment due to flaps and the lift at
which flow separation first occurred are
compared with experiment


NACA RM A51E29

LOW-SPEED CHARACTERISTICS OF A 450 SWEPT
WING WITH LEADING-EDGE INLETS. Robert E.
Dannenberg. August 1951. 48p. diagrs., photos.,
tab. (NACA RM A51E29)

Pressure-distribution and wake measurements were
made to determine the effect of leading-edge inlets
on the low-speed aerodynamic characteristics of a
450 swept wing that completely spanned the wind
tunnel. The inlets extended over approximately the
central third of the span of the wing and varied in
entrance height from 15 to 50 percent of the wing
thickness. The section characteristics were ob-
tained for a wide range of angles of attack and inlet-
velocity ratios.


NACA RM A51G31

THE FORCES AND PRESSURE DISTRIBUTION AT
SUBSONIC SPEEDS ON A PLANE WING HAVING 450
OF SWEEPBACK, AN ASPECT RATIO OF 3, AND A
TAPER RATIO OF 0.5. Carl D. Kolbe and
Frederick W. Boltz. October 1951. 159p. diagrs.,
photo., 22 tabs. (NACA RM A51G31)

Lift, drag, pitching-moment, and static -pressure
data for a model of a plane wing having an aspect
ratio of 3 and a taper ratio of 0.5 are reported. The
line joining the quarter-chord points of the airfoil
sections was swept back 450. The airfoil sections
were the NACA 64A010 in planes inclined 450 to the
plane of symmetry. Data are presented for Reynolds
numbers from 4,000,000 to 18,000,000 at a Mach
number of 0.25 and for Mach numbers from 0.08 to
0.96 at a Reynolds number of 4,000,000. Also in-
cluded are data at a Reynolds number of 8,000,000
for Mach numbers ol 0.08, 0.25, and 0.60. The in-
vestigation included a study of the effect of various
seals at the model-turntable juncture on the forces,
moments, and pressure distribution on the model.
Pressure data at seven spanwise stations on the
model are presented in tabular form.




NACA RM A51G31a

SUMMARY OF RESULTS OF A WIND-TUNNEL IN-
VESTIGATION OF NINE RELATED HORIZONTAL
TAILS. Jules B. Dods, Jr. and Bruce E. Tinling.
October 1951. 105p. diagrs.. photos., 2 tabs.
(NACA RM A51G31a)








a

A compilation of data is presented for models of nine
related horizontal tails. The majority of the results
were obtained at a Mash number of approximately
0.20. Three of the models were tested throughout the
subsonic Mach number range to a maximum of 0.94.
The Reynolds number range was from 2 to 4 million.
The models had aspect ratios from 2 to 6, angles of
sweepback from 5.70 to 450, and had 30-percent-
chord, sealed, plain flaps. The lift coefficient, hinge
moment coefficient, and pressure coefficients across
the elevator nose seal are presented. The effects of
sweepback, aspect ratio, and Mach number on the lift
and hinge-moment parameters are summarized.
Comparisons of the experimental results with theo-
retical calculations are presented.




NACA RM A52A14a

LOW-SPEED AERODYNAMIC CHARACTERISTICS
OF A LARGE-SCALE 600 SWEPT-BACK WING WITH
HIGH LIFT DEVICES. Mark W. Kelly. March 1952.
54p. dzagrs., photo., 7 tabs. (NACA RM A52A14a)


Tests were made in the Ames 40- by 80-foot wind
tunnel of a 600 sweptback wing with fuselage and of
the combination with various configurations of double-
slotted flaps, split flaps, leading-edge slats, and
ailerons. Lift, drag, pitching-moment, and rolling-
moment data are presented as well as pressure dis-
tribution data over the slat, the wing, and flaps. The
tests covered a Reynolds number range from 4.0x 106
to 10.0 x 106. The use of double-slotted flaps and
leading-edge slats extended the linear range of the
pitching-moment curve from a lift coefficient of about
0. 35 to a lift coefficient of about 0. 88.





NACA RM A52B19

LOW-SPEED AERODYNAMIC CHARACTERISTICS OF
A LARGE-SCALE 450 SWEPT-BACK WING WITH
PARTIAL-SPAN SLATS, DOUBLE-SLOTTED FLAPS,
AND AILERONS. Harry A. James. April 1952.
101p. diagrs., photos., 9 tabs. (NACA RM A52B19)


Tests were made in the Ames 40- by 80-foot wind
tunnel of a 450 sweptback wing with fuselage in com-
bination with various configurations of double-slotted
flaps, leading-edge slats, and ailerons Lift, drag,
pitching-moment, and rolling-moment data are pre-
sented as well as pressure-distribution data over the
slat, wing, and flaps. The lests covered a Reynolds
number range from 2.5 x 106 to 8 x 106. The use of
double-slotted flaps and leading-edge slats extended
the near linear range of the pitching moment from a
lift coefficient of about 0.76 to a lift coefficient of
about 1.41.


NACA
RESEARCH ABSTRACTS NO.56

NACA RM A52D01b

BOUNDARY-LAYER MEASUREMENTS ON SEVERAL
POROUS MATERIALS WITH SUCTION APPLIED.
George B. McCullough and Bruno J. Gambuccl. 'J-*
1952. 26p. diagrs., photos., tab. (NACA RM
A52D01b)

Boundary-layer velocity profiles were measured on
several porous materials for a stream velocity of
165 feet per second with ratios of suction velocity to
free-stream velocity from 0 to 0.0162. The measure-
ments were made on the wall of a small wind tunnel
in a region of nearly zero pressure gradient and in
which transition from laminar to turbulent flow was
occurring on the smooth impervious wall. The ef-
fects of surface roughness and suction velocity are
shown.




NACA RM A52JOB

DOWNWASH CHARACTERISTICS AND VORTEX-
SHEET SHAPE BEHIND A 630 SWEPT-BACK WING-
FUSELAGE COMBINATION AT REYNOLDS NUM-
BER OF 6. 1 x 106. William H. Tolhurst, Jr.
December 1952. 45p. diagram photo. (NACA
RM A52J08)

The downwash characteristics and vortex-sheet
shape have been experimentally determined for a 630
sweptback wing of aspect ratio 3. 5 with a fuselage
and with boundary-layer control on the wing to pre-
vent separation. Downwash-angle contour maps are
presented for nine vertical transverse planes located
between 0. 57 and 2. 71 semispans aft of the 0. 25
mean aerodynamic chord point for three angles of
attack at a Reynolds number of 6. 1 x 106. The shape
of the vortex sheet is presented for the various sur-
vey planes and angles of attack. Theoretical and
experimental downwash angles are compared.




NACA RM A53A14

EXPERIMENTAL AND THEORETICAL STUDY OF
THE INTERFERENCE AT LOW SPEED BETWEEN
SLENDER BODIES AND TRIANGULAR WINGS.
Edward J. Hopkins and Hubert C Carel. May 1953.
40p. diagrs., photos., tab. (NACA RM A53A14)


Measurements of the forces and moments were made
at a Mach number of 0.25 for several geometrically
similar triangular wings (aspect ratio, 2.0) in the
presence of a slender body of revolution and for the
wings combined with the body. The ratios of maxi-
mum body diameter to wing span were 0.196, 0.259,
0.343, and 0.500. Comparisons are made between
the experimental results and results calculated by
the method given in NACA RM A51G24 which employs
the theories of Weissinger, Multhopp, and Lennertz.








NACA
RESEARCH ABSTRACTS NO. 56




NACA RM E50A04

APPLICATION OF BLADE COOLING TO GArS TUR-
BINES. Herman H. Ellerbrockr, Jr. and Louis J.
Schaler, Jr. May31, 1950. 102p. diagram photor.
(NACA RM E50A4)


A review of current knowledge on turbine-blade cool-
ing and a description of pertinent NACA Investia-
tions are presented. The gain, which can be ob-
tained by operating turbojet and turbine-propeller
engines at high ga temperatures, are shown-
Analyses are presented, which give the performance
possibilities of various methods of blade cooking as
well as the possibility of using nonstrategic blade
materlae. Finally, the performance characteristics
of cooled turbines and of engineer using cooled tur-
blues are briefly discussed.




NACA RM 850AZ

PRELIMINARY ANALTRIS OF PROBLEM OF
DETERMINING EXPERIMENTAL PERFORMANCE
OF AIR-COOLED TURBINE. I METHODS FOR
DBTERMINING HEAT-TRANSFER CBARACTERB-
TICS. Herman R. Ellerbrock, Jr. and Robert R.
lemenrr. June 12,1950. 48p. dagrs. (NACA
RM E50AuS)


In determining the exprimental performance of an
air-caoled turbine, the heat-transfer characteristics
mrut be evaluated. The suggested formulas that are
required to determine these characteristics are pre-
seated. The formulas have a form in which depend-
ent parameters are expresd as unknown functions
of lauependent parameters Methods of eprimen~-
Ing to determine these functions ae suggested. In
some cases general heat-transfer discussion that
lead to te suggested forms of the formulas ae
givn.




NACA RM 850A06

PRELIMINARY ANALYSIS OF PROBLEM OF
DETERMINING EXPERIMENTAL PERFORMANCE
OF AIR-COOLED TURBINE. II: METHODS FOR
DETERMINING COOLING-AIR-FL~OW CHARACTER-
IBTICS. Herman H. Ellerbrock, Jr. June 7, 1950.
20p. diagr. (NACA RM E50A06)


In the determination of the perfo,.r.nee of an air-
cooled turbine, the cooling-air-flow characteristics
between the root and the tip of the blades must be
evaluated. The methods, which must be verified and
the unknown functions evaluated, that are expected to
permit the determination of pressure, temperature,
and velocity through the blade cooling-air passages
from specific investigations are presented-


NACA RM E50A31

ALTITUlDE-CHAMBER PERFORMANCE OF
BRITISH ROLISl-ROTCE NENE II ENGINE. II -
18.00-INCH-DIAMETER JET NOZZLE. Ralph E.
Grey, Virginia L. Brightwell and Zelmar Barson.
July 10, 1950. 60p. diagrs., 2tabs. (NACA
RMn E50A31)


An attitude-chamber investigation of British Rolla-
Royce Nene II turbojet engine was conducted over
range of altitudes from sea level to 65,000 feet and
ram pressure ratios from 1.10 to 3.50, using an
18.00-inch-diameter jet nozzle. The 18.00-inch-
diameter jet nozzle gave slightly lower values of
net-thrust specific fuel consumption than either the
18.41- or the standard 18.75-inch-diameter jet
nozzles at high flight speeds. At low flight speed,
the 18.41 -inch-diameter jet nozzle gave the lowest
values of net-thrust specific fuel consumption.



NACA RM E50B10

ALTITUDE-CHAMBER PERFORMANCE OF
BRITJISH ROLI8-ROYCE NENE II ENGINE. IV -
EFFECT OF OPERATIONAL VARIABLES ON
TEMPERATURE DISTRIBUTION AT COM6BUSTION-
CHAMBER OUTLETS. SIdney C. Runtley. July 3,
1950. 17p. diagrl., photo. (NACA RM 550810)


Temperature surveys were made at the combustion-
chamber outlets of a British Rolls-Royce Nene II
engine. The highest mean nozzle-vane and mean gas
temperatures were found to occur at a radine
approximately 75 percent of the nozzle-vane length
from the inner ring of the nozzle-vane assembly.
Variations In engine speed, jet-nozzle area, simu-
lated altitude, and simulated flight speed altered the
temperature level but did not materially affect the
pattern of radial temperature distribution.



NACA RM E50C15

EFFECTS OF INLET ICING ON PERFORMANCE OF
AXIAL-FLOW TUTRBOJET ENGINE IN NATURAL
ICING CONDITIONS. Loren W. Acker andKenneth S.
Kleinknecht. May25, 1950. 61p. diagra.,photosl.,
tab. (NACA RM E50C15)


A flight investigation in natural scing conditions was
conducted to determine the effect of inlet ice forma-
tions on the performance of axtal-flow turbojet en-
gines. The results are presented for icing conditions
ranging from a lquid-water content of 0.1 to 0.9 grmm
per cubic meter and watler-droplet size from 10 to 27
micron at amblent-air temperature from 130 to
26o F. The data show time histories of jet thrust,
air flow, tail-pipe temperature, compressor efficien-
cy, and icing parameters for each Icing encounter.
The effect of inlet-guide-vane Icing was isolated and
shown to account for approximately one-half the total
reduction in performance caused by inlet icing.








10

NACA RM E50C28

COMPARISON OF OUTSIDE-SURFACE HEAT-
TRANSFER COEFFICIENTS FOR CASCADES OF
TURBINE BLADES. James E. Hubbartt. July 17,
1950. 30p. diagre., tab. (NACA RM E50C28)


A comparison of available results from heat-transfer
investigations on cascades of turbine blades is pre-
sented using the Nusselt equation. The conventional
correlation procedure is modified by defining the
Reynolds number by the average of the velocities and
the pressures around the blades. The correlation of
the results from impulse blades was improved by
using the Reynolds number defined by the average
velocity and pressure. The final comparison indi-
cated that several variables, which possibly influence
heat transfer, should be investigated.


NACA RM ESD03a

HEAT-TRANSFER AND OPERATING CHARACTER-
ISTICS OF ALUMINUM FORCED-CONVECTION AND
STAINLESS-STEEL NATURAL-CONVECTION
WATER-COOLED SINGLE-STAGE TURBINES. John
C. Freche and A. J. Diaguila. June 30, 1950. 48p.
diagrs.. ohoto., 2 tabs. (NACA RM E50D03a)


Two water-cooled turbines were operated one with
blades and disk of aluminum alloy cooled by forced
convection and the other with blades and disk of stain-
less steel cooled by natural convection. Heat-
transfer data, coolant pumping power, and operational
data, such as gas temperatures, coolant-flow rates,
and blade temperatures, are presented. The heat-
transfer results for both turbines agreed with those
for static-cascade and stationary-tube investigations.
Coolant pumping losses were not excessive and are
further reducible by discharging the coolant at a
smaller radius.


NACA RM E50D25

ANALYTICAL INVESTIGATION OF FLOW AND
HEAT TRANSFER IN COOLANT PASSAGES OF
FREE-CONVECTION LIQUID-COOLED TURBINES.
E. R. G. Eckert and Thomas W. Jackson. July 18,
1950. 45p. diagrs. (NACA RM E50D25)


The theoretical aspect of the problem of free-
convection liquid cooling of turbine blades was Inves-
tigated using boundary-layer calculations. A method
of improving the heat transfer in a hole with a small
diameter-to-length ratio is described. Comparisons
of theoretical equations with experimental results for
a free-convection water-cooled turbine are made and
the effect of Coriolls forces on the cooling effect Is
discussed.


NACA RM E50D27

PERFORMANCE OF 24-INCH SUPERSONIC AXIAL-
FLOW COMPRESSOR IN AIR. II COMPRESSOR
PERFORMANCE WITH INLET GUIDE VANES.
Melvin J. Hartmann and Edward R. Tyal. July 10,
1950. 15p. diagrs. (NACA RM E50D27)


NACA
RESEARCH ABSTRACTS NO. 56


The use of inlet guide vanes with the 24-Inch super-
sonic compressor resulted. in a decrease in maxi-
mum pressure ratio and adiabatic efficiency,- ad a
slight increase in equivalent mass flow. Thi losatir
total pressure and efficiency resulted from reduced
diffusion in the rotor-blade passages, increased
shock losses at the higher entrance Mach number,
and increased boundary-layer thickness, separation,
and transfer of mass flow toward the rotor hub. The
unsteady flow field created at the compressor en-
trance by the guide-vane wakes is also responsible
for some of the losses. This inherent loss will be
encountered whenever inlet guide vanes are used
with the shock-in-rotor type of supersonic com-
pressor.






NACA RM E50E03

INVESTIGATION OF AERODYNAMIC AND ICING
CHARACTERISTICS OF WATER-INERTIA-
SEPARATION INLETS FOR TURBOJET ENGINES.
Uwe von Glahn and Robert E. Blatz. July 26,1950.
54p. diagrs., photos., 3 tabs. (NACA RM E50E03)


Aerodynamic and Icing Investigations of several In-
ternal water-inertia-separation inlets designed to
prevent entrance of water into a turbojet engine in an
icing condition are presented. Comparisons of total-
pressure loss, mass flow, and icing characteristics
are made. Complete icing protection of Inlet guide
vanes was not achieved. Approximately 8 percent of
the volume of water entering the Inlets remained in
the air. For nonicing operation, total-pressure
losses were comparable to those of direct-ram Inlets.
Under icing conditions, considerable total-pressure
losses were obtained with inertia-separation inlets.





NACA RM E50E04

NUMERICAL SOLUTION OF EQUATIONS FOR ONE-
DIMENSIONAL GAS FLOW IN ROTATING COOLANT
PASSAGES. W. Byron Brown and Richard J.
Rossbach. June 26, 1950. 119p. diagrs., tabs.
(NACA RM E50E04)


In order to analyze the compressible flow in a blade
coolant passage of an air-cooled gas turbine, a one-
dimensional analysis was made including the com-
bined effects of area change, wall friction, heat
transfer, and rotation. The necessary numerical
methods for the simultaneous solution of the result-
ing equations expressing the conservation of energy
and momentum in differential form are presented.
The solution of these equations can be greatly
simplified with little loss in accuracy by employing
constant mean values for the coefficients in the
energy equation. Tables are presented in order to
facilitate the required computations with the mini-
mum of interpolation.








NACA
RESEARCH ABSTRACTS NO. 56


NACA RM E50E18

PRELIMINARY ANALYSIS OF PROBLEM OF
DETERMINING EXPERIMENTAL PERFORMANCE
OF AIR-COOLED TURBINE. m METHODS FOR
DETERMINING POWER AND EFFICIENCY. Herman
H. Ellerbrock, Jr. and Robert R. Ziemer. August 2,
1950. 55p. diagrs. (NACA RM E50E18)


Suggested formulas are given for determining air-
cooled turbine-performance characteristics, such as
power and efficiency, as functions of certain param-
eters. These functions, generally being unknown,
are determined from experimental data obtained
from specific investigations. Special plotting
methods for isolating the effect of each parameter
are outlined.



NACA RM E50E22

PRELIMINARY ANALYSIS OF EFFECTS OF AIR
COOLING TURBINE BLADES ON TURBOJET-
ENGINE PERFORMANCE. Wilson B. Schramm,
Alfred 1. Nachtlgall and Vernon L. Arne. August 2,
1950. 34p. diagrs., tab. (NACA RM E5OE22)


The effects of turbine-blade cooling on engine per-
formance were analytically Investigated for a turbo-
jet engine in which cooling air is bled from the en-
gine compressor. The analysis was made for a con-
stant turbine-inlet temperature and a range of
altitudes to determine the minimum cooling require-
ments to permit substitution of nonstrategic mate-
rials in turbine blading. The results indicate that,
for a constant inlet temperature, air cooling of the
turbine blades increases the specific fuel consump-
tion and decreases the thrust of the engine. The
highest possible cooling effectiveness is desirable to
minimize coolant weight flow and its effects on engine
performance.




NACA RM E50FO2

EXTENSION OF BOUNDARY-LAYER HEAT-
TRANSFER THEORY TO COOLED TURBINE
BLADES. W. Byron Brown and Patrick L.
Donoughe. August 11, 1950. 51p. diagrs., tab.
(NACA RM E50F02)


An equation for average heat transfer of a surface
was derived when the boundary layer changed from
laminar to turbulent. Influences on the heat transfer
through a laminar boundary layer of Mach number,
temperature ratio (gas temperature divided by wall
temperature), and exponents of gas-property temper-
ature relations were shown to be relatively small for
air with Mach numbers less than 2 and temperature
ratios between I and 4. Good agreement was ob-
tained with experimental results from cylinders, an
airfoil, and turbine blades.


II

NACA RM E50F05

EXPERIMENTAL AND ANALYTICAL STUDY OF
BALANCED-DIAPHRAGM FUEL DISTRIBUTORS
FOR GAS-TURBINE ENGINES. David M. Straight
and Harold Gold. August 14, 1950. 62p. diagrs.,
photos. (NACA RM E50F05)


Eight possible fuel distributor arrangements for dis-
tributing fuel equally to a plurality of spray nozzles
are presented and analyzed with emphasis on meet-
ing fuel-system requirements of fuel-distribution
accuracy, spray-nozzle pressure variations, and
fuel-system pressures. The experimental perform-
ance of three distributors is discussed. Data are
presented for a fuel distributor-model distributing
19.5 to 862 pounds per hour of fuel to each of 10
spray nozzles with a maximum deviation of 3.3 per-
cent from equal distribution. Methods for distribut-
ing wider ranges of fuel-flow rates are included.


NACA RM E50F09

ANALYTICAL DETERMINATION OF LOCAL SUR-
FACE HEAT-TRANSFER COEFFICIENTS FOR
COOLED TURBINE BLADES FROM MEASURED
METAL TEMPERATURES. W. Byron Brown and
Jack B. Esgar. August 11, 1950. 66p. diagram.
(NACA RM E50F09)


Analytical methods are presented for the determina-
tion of local values of outside and inside heat-
transfer coefficients and effective gas temperatures
by use of turbine-blade-temperature measurements.
The methods are derived for a number of configura -
tions that can be applied to typical cooled-turbine-
blade shapes as well as to other types of heat-
transfer apparatus.

NACA RM E50F14

EFFECT OF RETRACTABLE IGNITION PLUG ON
PLUG FOULING BY CARBON DEPOSITS. Jerrold D.
Wear and Theodore E. Locke. August 24, 1950. 25p.
diagrs., photos. (NACA RM E50Fl4)


Investigations of ignition-plug fouling, starting, alti-
tude combustion efficiencies, altitude operational
limits, and combustor-outlet-temperature distribu-
tion with a retractable ignition plug and with a stand-
ard ignition plug were conducted. The retractable
ignition plug did not become fouled when investigated
with fuels and at engine conditions that did foul the
standard plug. Starting, altitude combustion effi-
ciencies, and altitude operational limits determined
with the standard plug were unaffected by use of the
retractable plug. A slightly improved temperature
distribution was obtained with the retractable plug.


NACA RM E50GI1

INVESTIGATION OF EFFECTS OF INLET-AIR
VELOCITY DISTORTION ON PERFORMANCE OF
TURBOJET ENGINE. E. William Conrad and Adam
E. Sobolewski. September 13, 1950. 41p. diagrs.,
photos. (NACA RM E50GI1l)








12



To determine effect of nonuniform inlet-air veloci-
ties a full scale, axial-flow turbojet engine was in-
vestigated in Lewis altitude wind tunnel at altitudes
from 20,000 to 50,000 feet, 0.21 flight Mach number
and corrected engine speeds from 77.3 percent of
rated speed to rated speed. Total-pressure varia-
tions as large as 103 pounds per square foot in
radial direction and 90 pounds per square foot in
circumferential direction at 30,000 feet were ob-
tained. With the distortions investigated, net thrust
varied between 0.95 and 1.03 of the thrust with uni-
form inlet-air distribution. Similarly the ratio of
specific fuel consumption varied from 1.00 to 1.04.
Within the range of this investigation, the effects of
nonuniform inlet velocity were not serious for the
engine investigated.





NACA RM E50014

COMBUSTION EFFICIENCIES IN HYDROCARBON-
AIR SYSTEMS AT REDUCED PRESSURES. Robert
R. Hibbard, Isadore L. Drell, Allen J. Metzler and
Adolph E. Spakowsld. September 13, 1950. 12p.
diagrs. (NACA RM E50G14)


In preliminary results obtained with quiescent fuel-
air mixtures and with small diffusion flames, com-
bustion efficiencies close to 100-percent were ob-
tained at pressures much lower than those found in
turbojet combustors at 60,000-foot altitude; in gener-
al, efficiencies were high at pressures approaching
the limiting values for inflammation. Apparently, it
is possible to burn flames with high efficiency at high
altitudes if sufficient volume Is available and un-
desired quenching processes do not occur. The
reaction-zone volume required to effect a given rate
of heat release rapidly increased with a decrease in
pressure. It appears, however, that only a small
fraction of the volume available in a turbojet com-
bustor Is required to effect the engine's design rate
of heat release.





NACA RM E50104

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADES IN TURBOJET ENGINE. I -
ROTOR BLADES WITH 10 TUBES IN COOLING-AIR
PASSAGES. Herman H. Ellerbrock, Jr. and
Francis S. Stepka. December 12, 1950. 76p.
diagrs., photos., 2 tabs. (NACA RM E50I04)


The experimental results of air cooling turbine
blades are presented herein. A method was devel-
oped for predicting cooled-blade temperatures by
use of a temperature-difference ratio defined as v
and a calculated value of the solid-uncooled-blade
temperature. By the use of this method predictions
were made of the allowable turbine-inlet tempera-
tures that may be attained with air cooling of turbine
blades.


NACA
RESEARCH ABSTRACTS NO.56


NACA RM E50114

EXPERIMENTAL INVESTIGATION OF AIR-COObED
TURBINE BLADES IN TURBOJET ENGINE. II -
ROTOR BLADES WITH 15 FINS IN COOLING-AIR
PASSAGES. Robert 0. Hickel and Herman H.
Ellerbrock, Jr. November 20, 1950. 56p. diagrs.,
photos., 3 tabs. (NACA RM E50114)


A turbine-blade-cooling investigation was conducted
on a modified production turbojet engine to determine
experimentally the cooling effectiveness of a cooled
blade with 15 fins Inserted in the hollow blade shell.
The results showed that the 15-fin blade cooled very
well at the mldchord, but that the leading and trailing
edges were considerably hotter. The 15-fin blade
showed Increased cooling effectiveness, particularly
near the midchord, over that of a blade with 10 tubes
inserted in the blade shell investigated earlier.





NACA RM E50115

ALTITUDE PERFORMANCE OF J35-A-17 TURBO-
JET ENGINE IN AN ALTITUDE CHAMBER. K. R.
Vincent and B. M. Gale. January 3, 1951. 53p.
diagrs., 3 tabs. (NACA RM E50I15)


Simulated flight performance of a J35-A-17 turbojet
engine in an attitude chamber is presented for range
of altitudes from 20,000 to 30,000 feet at flight Mach
number of 0.62 and for range of flight Mach numbers
from 0.42 to 1.22 at altitude of 30,000 feet. Per-
formance variables jet thrust, net thrust, air flow,
fuel flow, net-thrust specific fuel consumption, and
tail-pipe total gas temperature are presented as
functions of engine speed. Engine performance
generalized for altitudes up to 30,000 feet and for
flight Mach numbers at which critical flow exists in
exhaust nozzle with the exception of corrected net
thrust and net thrust specific fuel consumption.




NACA RM E50J06

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADES IN TURBOJET ENGINE. mI -
ROTOR BLADES WITH 34 STEEL TUBES IN
COOLING-AIR PASSAGES. Robert 0. HIckel and
Gordon T. Smith. December 11, 1950. 39p.
diagre., photos., 2 tabs. (NACA RM E50J06)


An experimental investigation to determine the
cooling effectiveness of air-cooled turbine blades
with 34 steel tubes Inserted in the blade shell was
made in a production turbojet engine over a range of
engine speeds from 4000 to 9000 rpm and coolant
flows from about 0.005 to 0.125 pound per second
per blade. Results showed that only a limited in-
crease, if any, in cooling effectiveness was obtained
with the 34-tube blade over that of the 10-tube and
15-fin blades, which were previously Investigated.


*i








NACA
RESEARCH ABSTRACTS NO. 56



NACA RM E54T23

DETERMINATION OF GAS-TO-BLADE CONVEC-
TION HEAT-TRANSFER COEFFICIENTS ON A
FORCED-CONVECTION, WATER-COOLED SINGLE-
BTAGE ALUMINUM TURBINE. John C. Freche and
Eugene F. Schum. Janury 3, 1951. 17p. diagre.,
tab. (NACA RM E5Q452)


Gma-to-blade convection heat-transfer coefficients
were obtained on a forced-convection, water-cooled,
single-stage alumlaum turbine over a range of gas-
flow Reynolds numbers, inlet-gas temperatures, and
at several computed gas-flow inlet angles. The con-
Tection coefficients were correlated by the conven-
Honal Nusselt equation. Individual correlation curves
were obtained for each gas-flow inlet angle. Com-
parison of the correlation curve obtained at the de-
sIgn ga-flow Inlet angle with static -cascade results
and with gas-to-blade heat-transfer result computed
by boundary-layer theory showed close agreement.




NACA RM E50K01

INVESTIGATIONS OF 81DT CONFIGURATIONS FOR
FILM-COOLED TURBINE BLADES BY FLOW
VIBUALIZATION METHODS. E. R. G. Eckert,
Thomas W. Jackson and Allen C. Francisco.
January IS, 191. 57p. diagrs., photos. (NACA
RM E50K01)


Two visual methods of obtaining qulalitative Informa-
tion on the effectiveness of different slot configura-
tions used to distribute cooling air around turbine
blades are discussed. One method indicates the
coolant coverage of the blade by means of a paint
which changes its color under the influence of ammo-
ala mixd with the cooling ar. The other method
makes the cooling air visible by mixing amoke with
It. Cooling filma generated by rowrs of cooking holes
flush to the blade surface, rows of holes ending in
grooves, and slata were studied-




NACA RML E50KIO

EXPERIMENTAL INVESTIGATION OF EFFECTS OF
DESIGN CHANGES ON PERFORMANCE OF LARGE-
CAPACITY CENTRIFUGAL COMPRESSORS. Joseph
R. Withee, Jr., Karl KOvach and Ambrose Ginsburg.
March 2, 1951. 35p diagrs., photos., tab. (NACA
RM E50K10)


An investigaton wna conducted to determine the of -
fects of several design changes on the performance
of large-capacity, double-entry, centrifugal com-
preasors. Delga changes that resulted in major im-
provements in performance were the ue of a fuly
machined parbablic-blade inducer, increased im-
peller Inlet-to-tip radina ratio, and a fully machined
diffuser.


NACA RM E501L15

SURVEY OF ADVANTAGES AND PROBLEMS
ASSOCIATED WITH TRANSPIRATION COOLING
AND FILM COOLING OF GAS-TURBINE BLADES.
E. R. G. Eckrert and Jackr B. Eagar. February 12,
1951. 39p. diagrs. (NACA RM E50Kl5)


An introduction to transpiration and film cooling
methods as applied to gas-turbine blades is pre-
sented. The physical processes of these cooling
methods are expnlaned and information available for
prediction of blade temperatures and beat-transfer
rates la surveyed.





NACA RM E50LOT

THERMAL CONDUCTIVITY OF 14 METAIAI AND
ALLOYS UP TO 1100oF. Jerry E. Evane, Jr.
March, 1951. 15p. diagra., tab. (NACA
RM E50LO't)


The thermal conductivity of 14 metals and alloys was
determined in temperature ranges having a maximum
of 11000 F. The metals tested include steels, high-
temperature alloys, molybdenum dlialicide, alumi-
num alloys, brass, and silver. A comparison
method was used to obtain the results in which the
thermal conductivity of the test sample was com-
pared with the conductivity of high-purity lead.






NACA RM E50L20

AVERAGE OUTSIDE-SURFACE HEAT-TRANSFER
COEFFICIENTS AND VELOCITY DISTRIBUTIONS
FOR HEATED AND COOLED IMPUILiE TURBINE
BLADES IN STATIC CASCADES. James E.
Hubbartt and Eugene F. Schum. March 9, 1951.
35~p. diagram photos. (NACA RMd E50L201


Gas-to-blade and blade-to-gas convective heat-
transfer coefficients were exprimentally obtained
from a static cascade of impulse blades that could
be heated or cooled. Blade peripheral velocities
were measured and compared with calculated veloc-
itles obtained from a theory derived for impulse-
type blades. The convective coefficients were cor-
related by the conventional Nuagelt equation wherein
Reynolds number was multiplied by a gas-to-blade
temperature ratio. Blade peripheral calculated
velocities were approximately 20 percent lower than
measured volelocte. Comparison of correlated
curves for the heated and cooled blade with heat-
transfer result computed by boundary-layer theory
showed agreemet within 4 percent at high expert-
mental values of Raynolds number and 23 percent at
low values.







14




NACA RM E51AO3

A SUMMARY OF DESIGN INFORMATION FOR
WATER-COOLED TURBINES. John C. Freche.
March 9, 1951. 26p. diagram. (NACA RM E51A03)


Information pertinent to the design of water-cooled
turbines is tabulated. A design procedure is out-
lined. Calculation methods, design, fabrication,
material, and operating considerations required for
water-cooled turbine design are discussed.



NACA RM E51Al9

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADES IN TURBOJET ENGINE. IV -
EFFECTS OF SPECIAL LEADING- AND TRAILING-
EDGE MODIFICATIONS ON BLADE TEMPERA-
TURE. Herman H. Ellerbrock, Jr., Charles F.
Zalabak and Gordon T. Smith. April 13, 1951.
71p. diagrs., photos., 3 tabs. (NACA RM E51A19)


Six air-cooled blade configurations with special
modifications for cooling the leading and trailing
edges were fabricated and experimentally investi-
gated in a production turbojet engine. Results of
this investigation are presented and compared with
results obtained from blade configurations previously
investigated. The most promising of these blade
designs may operate satisfactorily when fabricated
of nonstrategic material at present turbine-inlet gas
temperatures. General considerations necessary
for cooled-blade design are discussed.



NACA RM E51A22

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADES IN TURBOJET ENGINE. V -
ROTOR BLADES WITH SPLIT TRAILING EDGES.
Gordon T. Smith and Robert 0. Hickel. April 2,
1951. 17p. diagrs., tab. (NACA RM E51A22)


An air-cooled turbine-blade configuration with
special provision for cooling the trailing edge by
passing cooling air through a series of radial slots
was investigated. The results of this investigation
are presented and compared to the results obtained
from other blade configurations previously investi-
gated.





NACA RM E51CO9

EFFECT OF BLADE-SURFACE FINISH ON PER-
FORMANCE OF A SINGLE-STAGE AXIAL-FLOW
COMPRESSOR. Jason J. Moses and George K.
Serovy. April 16, 1951. 25p. diagrs., photos., 2
tabs. (NACA RM E51C09)


NACA
RESEARCH ABSTRACTS NO. 56




A set of modified NACA 5509-34 rotor and stated
blades was investigated with rough-machined, hahd -
filed, and highly polished surface finishes over a
range of weight flows at six equivalent tip speeds
from 672 to 1092 feet per second to determine the ef-
fect of blade-surface finish on the performance of a
single-stage axial-flow compressor. Surface-finish
effects decreased with increasing compressor speed
and with decreasing flow at a given speed. In general,
finishing blade surfaces below the roughness that may
be considered aerodynamically smooth on the basis of
an admissible-roughness formula will have no effect
on compressor performance.


NACA RM F51C29

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADES IN TURBOJET ENGINE. VI -
CONDUCTION AND FILM COOLING OF LEADING
AND TRAILING EDGES OF ROTOR BLADES.
Vernon L. Arne and Jack B. Esgar. May 18, 1951.
51p. diagrs., 3 tabs. (NACA RM E51C29)


A turbine-blade-cooling investigation was conducted
on a modified commercial turbojet engine to deter-
mine experimentally the cooling and pressure-drop
characteristics of a copper-clad conduction-cooled
blade and three variations of a film-cooled blade.
The copper-clad blade had very low temperature
gradients combined with structural rigidity. AUll the
film-cooled blades had high leading- and trailing-
edge cooling effectiveness at high cooling-air flows,
but the effectiveness tended to decrease rapidly as
cooling-air flow was reduced below some critical
value.


NACA RM E51D12


PRELIMINARY STUDY OF CIRCULATION IN AN
APPARATUS SUITABLE FOR DETERMINING
CORROSIVE EFFECTS OF HOT FLOWING LIQUIDS.
Leland G. Desmon and Don R. Mosher. June 1951.
17p. diagrs., photos. (NACA RM E51D12)


A simple apparatus for inducing flow of a liquid me-
dium in toroidal shaped channels is described. No
pumps, valves or flow meters are required. The ap-
parhtus is particularly applicable to the determina-
tion of the corrosive effects of flowing liquid metals
on structural materials. Preliminary tests, wherein
fluid velocities of 25 feet per second were obtained,
are described.




NACA RM E51E07

SOME EFFECTS OF SOLIDITY ON TURNING
THROUGH CONSTANT-THICKNESS CIRCULAR-ARC
GUIDE VANES IN AXIAL ANNULAR FLOW. Harry
Mankuta and Donald C. Guentert. August 1951. 19p.
diagrs. (NACA RM E51E07)









NACA
RESEARCH ABSTRACTS NO. 56



An investigation was conducted on sheet-metal, cir-
cular arc compressor inlet guide vanes in an annular
cascade with untapered walls to determine the effect
of solidity on turning through a blade row. Guide
vanes of 300 and 400 camber were investigated over
a range of solidity from 0.5 to 4.0. An equation
similar in form to Constant's rule that may be used
to predict turning angles in cascades of configuration
similar to that of this investigation was obtained from
the data. Surveys made downstream of the guide
vanes indicated that lower solidities than those pres-
ently used may be employed without exceeding load-
ing limitations.



NACA RM E51E18

DETERMINATION OF BLADE-TO-COOLANT
HEAT-TRANSFER COEFFICIENTS ON A FORCED-
CONVECTION, WATER-COOLED, SINGLE-STAGE
TURBINE. John C. Freche and Eugene F. Schum.
July 1951. 25p. diagrs., 2 tabs. (NACA RM
E51E18)

Blade-to-coolant convective heat-transfer coeffi-
cients were obtained on a forced-convection water-
cooled single-stage turbine over a large laminar
flow range and over a portion of the transition range
between laminar and turbulent flow. The convective
coefficients were correlated by the general relation
for forced-convection heat transfer with laminar
flow. Natural-convection heat transfer was negligi-
ble for this turbine over the Grashof number range
investigated. Comparison of turbine data with sta-
tionary tube data for the laminar flow of heated
liquids showed good agreement. Calculated
average midspan blade temperatures using theo-
retical gas-to-blade coefficients and blade-to-
coolant coefficients from stationary-tube data re-
sulted in close agreement with experimental data.




NACA RM E51E21

EVALUATION OF CENTRIFUGAL COMPRESSOR
PERFORM ANCE WITH WATER INJECTION.
William L. Beede, Joseph T. Hamrick and Joseph R.
Withee, Jr. July 1951. 14p. diagrs. (NACA RM
E51E21)

The effects of water injection on a compressor are
presented. To determine the effects of varying
water-air ratio, the compressor was operated at a
constant equivalent impeller speed over a range of
water-air ratios and weight flows. Operation over a
range of weight flows at one water-air ratio and two
inlet air temperatures was carried out to obtain an
indication of the effects of varying inlet air temper -
ature. Beyond a water-air ratio of 0.03 there was
no increase in maximum air-weight flow, a negli-
gible rise in peak total-pressure ratio, and a de-
crease in peak adiabatic efficiency. An increase in
inlet air temperature resulted in an increase
in the magnitude of evaporation. An analysis of data
indicated that the magnitude of evaporation within
the compressor impeller was small.


15


NACA RM E51E23

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADES IN TURBOJET ENGINE. VII -
ROTOR-BLADE FABRICATION PROCEDURES.
Roger A. Long and Jack B. Esgar. September 1951.
30p. diagrs., photos. (NACA RM E51E23)


Fabrication procedures for a variety of air-cooled
turbine-blade configurations, which were used in
experimental investigations to determine the effec-
tiveness of air-cooled blades in a turbojet engine,
are presented. The blade shells were cast or
formed from sheet metal or tubes. Additional heat-
transfer surface was provided in the coolant passage
by brazing fins or small tubes to the blade shell.
Fabrication of blades with special methods of cooling
leading and trailing edges is also discussed.



NACA RM E51E24

CALCULATED EFFECTS OF TURBINE ROTOR-
BLADE COOLING-AIR FLOW, ALTITUDE, AND
COMPRESSOR BLEED POINT ON PERFORMANCE
OF A TURBOJET ENGINE. Vernon L. Arne and
Alfred J. Nachligall. August 1951. 24p. diagrs.,
2 tabs. (NACA R.M E51E241

Effects of air-cooling turbine rotor blades on per-
iormance of a turbojet engine were calculated for a
range of altitudes frorn sea level to 40,000 feet and
a range of coolant flows up to 3 percent of compres-
sor air flow, for two conditions of coolant bleed
from the compressor. Bleeding at required coolant
pressure resulted in a sea-level thrust reduction
approximately twice the percentage coolant flow and
in an increase in specific fuel consumption approxi-
mately equal to percentage coolant flow. For any
fixed value of coolant flow ratio the percentage
thrust reduction and percentage increa e in
specific fuel consumption decreased with altitude.
Bleeding coolant at the compressor discharge re-
sulted in an additional 1 percent loss in performance
at sea level and in smaller increase in loss of per-
formance at higher altitudes.



NACA RM E51FO4

ANALYTICAL INVESTIGATION OF TWO LIQUID
COOLING SYSTEMS FOR TURBINE BLADES.
Thomas W. Jackson and John N. B. Livingood.
August 1951. 27p. diagrs. (NACA RM E51 F04)


A simplified analysis was made to determine flow
characteristics and heat transfer in the laminar and
turbulent flow regions for two turbine liquid-cooling
systems, a straight -through-flow system and a re-
circulatory, or loop, flow system. Nondimensional
charts, which simplify calculations of heat-transfer
coefficients and coolant temperatures, are pre-
sented. Results of a sample calculation showed
that, for a blade temperature oi 10000 F, an in-
crease in effective gas temperature of 3500 F could
be obtained by use of the loop cooling system instead
of the straight-through system.








16

NACA RM E51FO6

PERFORMANCE OF SINGLE-STAGE COMPRESSOR
DESIGNED ON BASIS OF CONSTANT TOTAL
ENTHALPY WITH SYMMETRICAL VELOCITY
DIAGRAM AT ALL RADHI AND VELOCITY RATIO
OF 0.7 AT ROTOR HUB. Jack R. Burtt and
Robert J. Jackson. September 1951. 14p. diagrs.
(NACA RM E51F06)

A typical inlet axial-flow compressor inlet stage,
which was designed on the basis of constant total
enthalpy with symmetrical velocity diagram at all
radii, was investigated. At a tip speed of 1126 feet
per second, a peak pressure ratio of 1.28 was ob-
tained at an efficiency of 0.76. At this tip speed, the
highest practical flow was 28 pounds per second per
square foot frontal area with an efficiency of 0.78.
Data for a rotor relative inlet Mach number range of
from 0.5 to 0.875 indicates that the critical value for
any stage radial element is approximately 0.80
for the stage investigated.



NACA RM E51F22

CALCULATIONS OF LAMINAR HEAT TRANSFER
AROUND CYLINDERS OF ARBITRARY CROSS
SECTION AND TRANSPIRATION-COOLED WALLS
WITH APPLICATION TO TURBINE BLADE COOL -
ING. E. R. G. Eckert and John N. B. Livingood.
September 1951. 57p. diagrs., 2 tabs. (NACA
RM E51F22)

An approximate method for development of flow and
thermal boundary layers in laminar regime on cylin-
ders with arbitrary cross section and transpiration-
cooled walls is obtained by use of Karman's integrat-
ed momentum equation and an analogous beat-flow
equation. Incompressible flow with constant property
values throughout boundary layer is assumed. Shape
parameters for approximated velocity and tempera-
tare profiles and functions necessary for solution of
boundary-layer equations are presented as charts,
reducing calculations to a minimum. The
method is applied to determine local heat-transfer
coefficients and surface temperatures in the laminar
region of transpiration-cooled turbine blades for a
given flow rate. Coolant flow distributions necessary
for maintaining uniform blade temperatures are also
determined.

NACA RM E51G10O

DETERMINATION AND USE OF THE LOCAL
RECOVERY FACTOR FOR CALCULATING THE
EFFECTIVE GAS TEMPERATURE FOR TURBINE
BLADES. Jack B. Esgar and Alfred L. Lea.
September 1951. 30p. diagrs., photos. (NACA RM
E51G10)

In an experimental investigation aof local recovery
factors for a blade having a pressure distribution
similar to that of a typical reaction-type turbine
blade, it was found that the recovery factors were
essentially independent of Mach number, Reynolds
number, pressure gradient, and position on the blade
surface except for regions where the boundary layer
was probably in the transition range from laminar to
turbulent. The recommended value of local subsonic
recovery factor for use in calculating the effective
gas temperature for gas turbine blades was 0.89.


NACA
RESEARCH ABSTRACTS NO. 56
NACA RM E51G30

PERFORMANCE OF A CASCADE IN AN ANNULAR
VORTEX-GENERATING TUNNEL OVER RANG&OF
REYNOLDS NUMBERS. Sidney Thurston and Ralph" .
E. Brunk. September 1951. 33p. diagrs. (NACA
RM E51G30)

Total-pressure deficiency for an annular cascade aof
65-(12)10 blades was measured at three radial sta-
tions over a range of Reynolds numbers from 50,000
to 250,000 and at angles of attack of 150 and 250.
The variation of turning angle and shape of static
pressure distribution at these stations is also shown.


NACA RM E51G31

PRELIMINARY INVESTIGATION OF MOLYBDENUM
DISULFIDE AIR-MIST LUBRICATION FOR
ROLLER BEARINGS OPERATING TO DN VALUES
OF I x 106 AND BALL BEARINGS OPERATING TO
TEMPERATURES OF 10000 F. E. F. Macks, Z. N.
Nemeth and W. J. Anderson. October 1951. 38p.
diagram photos., 2 tabs. (NACA RM E51G31)


The effectiveness of molybdenum disulfide MoB2 as a
bearing lubricant was determined at high temperature
and at high speeds. A 1-inch-bore ball bearing op-
erated at temperatures to 10000 F, a speed of 1725
rpm, and a thrust load of 20 pounds when lubricated
only with MoS2-air mist. A 75-millimeter-bore
cageless roller bearing, provided with a MoS2-syrup
coating before operation and lubricated with MoS2-air
mist during operation, operated at DN values to
I x 106 with a load of 368 pounds.

NACA RM E51H14

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADE IN TURBOJET ENGINE. VIII -
ROTOR BLADES WITH CAPPED LEADING EDGES.
Gordon T. Smith and Robert 0. Hickel. October
1951. 23p. diagrs., tab. (NACA RM E51H14)


An air-cooled turbine blade with a capped leading
edge was investigated in a modified commercial
turbojet engine over a range of engine speed from
4000 to 11,350 rpm. The cooling performance of the
capped-leading-edge configuration was superior to
all leading-edge cooling modifications previously in-
vestigated.


NACA RM E51Ill

INVESTIGATIONS OF AIR-COOLED TURBINE
ROTORS FOR TURBOJET ENGINES. I EXPERI-
MENTAL DISK TEMPERATURE DISTRIBUTION IN
MODIFIED J33 SPLIT-DISK ROTOR AT SPEEDS UP
TO 6000 RPM. Wilson B. Schramm and Robert R.
Ziemer. January 1952. 37p. diagrs., photos.,
2 tabs. (NACA RM E51Ill)

The experimental results of the disk temperature
distribution in an air-cooled J33 split-disk turbine
rotor fitted with nontwisted hollow air-cooled blades
with nine tubes in the coolant passage are presented.
These results were obtained at speeds up to 6000 rpm
and the data correlated so that the disk temperatures
at rated engine conditions could be estimated.







NACA
RESEARCH ABSTRACTS NO. 56


NACA M E5sU17

BLADE-TO-COOLANT HEAT-TRAMUFER RESULTS
AND OPERATING DATA FROM A NATURAL-
CONVECTION WATER-COOLED SINGLE-STAGE
TURBINE. Anthony J. Diagulla and John C. Freche.
November 1951. 22p. diagrs., tab. (NACA EM
Z51117)

Blade-to-coolant heat-transfer data and operating
data were obtained with a natural-convection water-
cooled turbine over a range of turbine speeds and
inlet-gas temperatures. The convective coefflcients
were correlated by the general relation for natural-
convection heal transfer. The turbine data were dis-
placed from a theoretical equation for natural con-
vection beat transfer in the turbulent region and from
natural-convection data obtained with vertical cylin-
ders and plates; possible disruption of natural con-
vection circulation within the blade coolant passages
was thus indicated. Comparison of nonlimensianal
temperature-ratio parameters for the blade leading
edge, midchord, and trailing edge indicated that the
blade cooling effectiveness is greatest at the mid-
chord and least at the trailing edge.




NACA RM E51J03

INVESTIGATIONS OF AIR-COOLED TURBINE
ROTORS FOR TURBOJET ENGINES. II MECHANI-
CAL DESIGN, STRESS ANALYSIS, AND BURST TEST
OF MODIFIED J33 SPLIT-DISK ROTOR. Richard H.
Kemp and Merland L. Moseson. January 1952. 46p.
diagrs., photos., tab. (NACA RM E51JO3)


A full-scale J33 air-cooled split turbine rotor was
designed and spin-pit tested to destruction. Stress
analysis and spin-pit results indicated that the rotor
was structurally sound. Operation of a similar rotor
in a J33 turbojet engine, however, showed that the
rear disk of the rotor operated at temperatures sub-
stantially higher than the forward disk. An extension
of the stress analysis to include the temperature
difference between the two disks indicated that engine
modifications are required to permit operation of the
two disks at more nearly the same temperature level.





NACARM E51JI10

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADES IN TURBOJET ENGINE. IX -
EVALUATION OF THE DURABILITY OF NONCRIT-
ICAL ROTOR BLADES IN ENGINE OPERATION.
Francis S. Stepka and Robert 0. Hickel. December
1951. 26p. diagrs., photos., 4 tabs. (NACA RM
E51J10)

The durability of five different structural or cooling
configurations or combination of both of air-cooled
blades made of noncritical materials was investi-
gated in a modified turbojet engine. The greater
part of the investigation was conducted at an engine
speed of 11,500 rpm, a turbine-inlet temperature of


17


approximately 16700 F, and a cooling-air to combus-
tion gas flow ratio per blade of 0.05. The results of
the investigation indicated that air-cooled blades
made of noncritical metals can be operated for ex-
tended periods of time in engines at current Inlet
temperatures; however, before these blades are con-
sidered completely satisfactory for gas-turbine ap-
plication, some means such as coatings is required
to inhibit the oxidation of the blades.




NACA RM E51KO2

EXPERIMENTAL INVESTIGATION OF COOLANT-
FLOW CHARACTERISTICS OF A SINTERED POROUS
TURBINE BLADE. Edward R. Bartoo, Louis J.
Schafer, Jr. and Hadley T. Richards. February
1952. 28p. diagrs., photos. (NACA RM E51K02)


Local cooling-air flow rates through the walls of a
sintered porous-metal turbine blade were measured
at room temperature for a range of pressure drops.
In order to check the validity of the correlation pro-
cedure, this procedure was used to correlate air-flow
rates through two porous disks at temperatures up to
6000 F. Data indicate the method permits room-
temperature flow data to be used for heat-transfer
work at elevated temperatures with reasonable accu-
racy. Cooling-air flow distribution around the pe-
riphery of the test blade Is presented for two Internal
cooling air pressures with the blade in a 10000 F gas
stream at a cascade-inlet Mach number of 0. 45.



NACA RM E51K08

EXPERIMENTAL INVESTIGATION OF THE HEAT-
TRANSFER CHARACTERISTICS OF AN AIR-
COOLED SINTERED POROUS TURBINE BLADE.
Louis J Schater, Jr., Edward R. Bartoo and Hadley
T Richards February 1952. 33p. diagrs.. photo.
(NACA RM E51KO8)


The results of a heat-transfer investigation on a
transpiration-cooled turbine blade are presented.
The investigation was made on a static cascade.
Blade-peripheral wall-temperature distributions
were measured and compared with the temperature
distribution on a convection-cooled blade. A cor-
relation of the blade-wall temperatures obtained at
various gas temperatures and gas weight flows was
made on the basis of a temperature-difference ratio.
Blade-wall temperatures calculated using an approxi-
mate theory for turbulent flow are compared witn the
measured blade-wall temperatures.



NACA RM E52AO4

OXIDATION-RESISTANCE MECHANISM AND OTHER
PROPERTIES OF MOLYBDENUM DISILICIDE. W A
Maxwell. March 1952. 17p. diagrs., photos., tab.
(NACA RM E52AO4)








18

The outstanding oxidation resistance of molybdenum
disilicide at 24000 F and above was found to depend
on the formation of a protective siliceous coating in
which a-cristobalite has been identified. Molybdenum
disilicide is not inherently resistant to oxidation and
in powdered form burns at low temperatures. Melt-
ing and casting experiments have demonstrated the
decomposition of the material at the melting point.
The room-temperature modulus of elasticity has been
determined and electric-resistivity data are given
to 20000 F

NACA RM E52B13

EXPERIMENTAL INVESTIGATION OF AIR-COOLED
TURBINE BLADES IN TURBOJET ENGINE. X -
ENDURANCE EVALUATION OF SEVERAL TUBE-
FILLED ROTOR BLADES. Jack B. Esgar and John
L. Clure. May 1952. 45p. diagrs., photos., 5 tabs.
(NACA RM E52B13)

Results are presented of endurance investigations of
a variety of air-cooled tube-filled turbine-rotor
blades to evaluate (1) new shell-to-base attachment
methods, (2) coatings for providing oxidation resist-
ance, and (3) durability of improved type of blade at
reduced coolant-flow rates. The improved type of
blade could operate satisfactorily at rated-speed
endurance conditions at coolant-flow ratios as low
as 0.03, but oxidation limited the blade life even
though the blades were protected with the coatings
that were investigated.

NACA RM E52C12

INVESTIGATIONS OF AIR-COOLED TURBINE
ROTORS FOR TURBOJET ENGINES. I- EXPERI-
MENTAL COOLING-AIR IMPELLER PERFORM-
ANCE AND TURBINE ROTOR TEMPERATURES IN
MODIFIED J33 SPLIT-DISK ROTOR UP TO SPEEDS
OF 10,000 RPM. Alfred J. Nachtigall, Charles F.
Zalabak and Robert R. Ziemer. May 1952. 42p.
diagrs., photos., 3 tabs. INACA RM E52C121


The experimental results of the cooling-air impeller
performance and the turbine rotor temperature dis-
tributions in an air-cooled J33 split-disk turbine
rotor fitted with air-cooled, tube-filled blades are
presented for engine speeds up to 10,000 rpm. Engine
alterations which reduced the temperature-level dif-
ference between the forward and rear disks of the
rotor permitted operation over the extended speed
range. Disk temperature data were correlated so
that disk temperatures at rated engine conditions
could be estimated. Calculated disk stresses at
raled engine conditions are presented and indicate
that the configuration investigated is structurally
sound


NACA RM E52D01

PRESSURE DROP IN COOLANT PASSAGES OF TWO
AIR-COOLED TURBINE-BLADE CONFIGURATIONS.
W. Byron Brown and Henry 0. Slone. June 1952. 57p.
diagrs., photos., tab. (NACA RM E52D01)


NACA
RESEARCH ABSTRACTS NO. 56

Friction coefficients were experimentally determined
for both isothermal and nonisothermal flows in a pres-
sure drop investigation conducted on two air-cooled
turbine-blade configurations. A reasonable sheck
was obtained between the air-cooled-blade frictiaq
coefficients. The data were used to develop a sim-'
plified flow equation for calculating over-all pressure
drops through air-cooled blades and to examine the
magnitude of entrance losses which might be expected
in the entrance sections of air-cooled blades. A pro-
cedure that may be used for the calculation of air-
cooled-blade pressure drops is presented.


NACA RM E52D21

COMPARISON OF CALCULATED AND EXPERIMEN-
TAL TEMPERATURES OF WATER-COOLED TUR-
BINE BLADES. Eugene F. Schum, John C. Freche
and William J. Stelpflug. July 1952. 36p. diagrs.,
3 tabs. (NACA RM E52D21)

Analytical methods were applied to calculate average
and local blade temperatures and maximum allowable
inlet gas temperatures for a forced-convection,
water-cooled aluminum turbine. Comparison of cal-
culated with experimental temperatures, obtained
over a gas-temperature range from 4000 to 16000 F,
resulted in generally good agreement. A stationary
water-cooled low-conductivity material (stainless
steel) blade with a high-conducitivty material
(copper) inserted in the trailing edge was also inves-
tigated over a gas-temperature range from 4000 to
9000 F and proved to be an effective method of reduc-
ing the trailing-edge temperature. Calculated trail-
ing edge temperatures agreed well with experimen-
tal data.


NACA RM E52F23

RADIANT HEAT TRANSFER FROM FLAMES IN A
SINGLE TUBULAR TURBOJET COMBUSTOR.
Leonard Topper. August 1952. 30p. diagrs., tab.
(NACA RM E52F23)

An experimental investigation of thermal radiation
from the flame of a single tubular turbojet-engine
combustor to the combustor liner is presented. The
effects of combustor inlet-air pressure, air mass
flow, and fuel-air ratio on the radiant intensity and
on the temperature and emissivity of the flame are
reported. The total radiation of the "luminous"
flames (containing incandescent soot particles) was
much greater (4 to 21 times) than the "nonluminous"
molecular radiation. The intensity of radiation from
the flame increased rapidly with an Increase in com-
bustor inlet-air pressure; it was affected to a lesser
degree by variations in fuel-air ratio and air mass
flow.



NACA RM E52F30

EXPERIMENTAL INVESTIGATION OF FREE-
CONVECTION HEAT TRANSFER IN VERTICAL
TUBE AT LARGE GRASHOF NUMBERS. E. R. G.
Eckert and A. J. Diaguila. August 1952. 37p.
diagras photos., tab. (NACA RM E52F30)








NACA
RESEARCH ABSTRACTS NO. 56

Local free-convection heat-transfer coefficients and
temperature fields in the turbulent flow range were
obtained within a vertical, stationary tube closed at
the bottom, heated along its walls, and having a
length-to-diameter ratio of 5. Convective heat-
transfer coefficients were correlated by the general
relations for free-convection heat transfer. These
coefficients, converted to dimensionless Nusselt
numbers were 35 percent below known relations for
vertical flat plates. Air temperature measurements
within the tube indicated a thin boundary layer along
the heated wall surface and unstable conditions in the
air flow.


NACA RM E52GII

CARBON-DEPOSITION CHARACTERISTICS OF
MIL-F-5624A FUELS CONTAINING HIGH-BOILING
AROMATIC HYDROCARBONS. Edmund R. Jlonash,
Jerrold D. Wear and William P. Cook. August
1952. 1 Ip. diagrs., photos., 2 tabs. (NACA
RM E52G11)

The effects of additions of typical high-boiling single-
ring and dicyclic aromatic hydrocarbons (tri-
isopropylbenzene and monomethylnaphthalene) on the
carbon deposition characteristics of MIL-F-5624A,
grade JP-3 and JP-4, fuel were determined in a
single J33 combustor. The addition of the aromatic
hydrocarbon components increased the carbon-
deposition characteristics of the base fuels. These
effects were satisfactorily predicted by the NACA K
factor and the smoking e'ienieicy correlations previ-
ously developed.



NACA RM E52H05

INVESTIGATION OF TITANIUM CARBIDE BASE
CERAMALS CONTAINING EITHER NICKEL OR CO-
BALT FOR USE AS GAS-TURBINE BLADES. C. A.
Hoffman and A. L. Cooper. August 1952. 33p.
photos., diagrs., 5 tabs. (NACA RM E52H05)


The following two materials were investigated for use
as gas-turbine-blade material: 65 percent TIC plus
20 percent Co plus 15 percent (CbTaTi)C and 65 per-
cent TiC plus 20 percent Ni plus 15 percent
(CbTaTi)C. Concurrently, the effectiveness of a
number of methods of preventing ceramal-blade-root
failure was studied. Blade temperatures from 15000
to 19000 F and wheel speeds from 10, 000 to 26, 000
rpm were used. Stellite 21 and S-816 alloy blades
were used for comparison. Prior to the blade eval-
uation, physical-prooerty evaluations of the ceramals
were made. The results indicate that both the 66. 5
percent TIC plus 18. 5 percent Co plus 15 percent
(CbTaTi)C and 66. 5 percent TiC plus 18. 5 percent
Ni and 15 percent (CbTaTi)C ceramals should be con-
sidered further for gas-turbine-blade use. The oxi-
dation resistance of either of these materials appeared
to be adequate for practical blade application. Fur-
ther work on blade root design and mounting methods
iV required to eliminate root failure as the factor
limiting the usefulness of these ceramals as gas tur-
bine blade material.


NACA RM E52H19


INVESTIGATION OF EFFECTS OF REYNOLDS
NUMBER ON LARGE DOUBLE-ENTRY CENTRIF-
UGAL COMPRESSOR Karl Kovach and Joseph R.
Withee. Jr. October 1952. 26p. diagrs. (NACA
RM E52H1119)

An investigation was conducted on a large-capacity
double-entry centrifugal compressor to determine the
effects of Reynolds number on over-all performance
and also to obtain some idea of the origin ana nature
of the effects. The decreases in over-all perform-
ance at design tip speed (1545 ft sec) for a change in
pressure altitude from 20, 000 to 50, 000 feet were:
(1) peak total-pressure ratio, 3. 4 percent, (2) maxi-
mum equivalent weight flow, 3. 8 percent, and (3)
peak efficiency, 0.02. The origin of at least part of
the Reynolds number effects in this compressor is in
the inlet and the inducer sections.

NACA RM E52H28

INVESTIGATION OF AXIALLY SYMMETRIC AND
TWO-DIMENSIONAL MULTINOZZLES FOR PRO-
DUCING SUPERSONIC STREAMS. Eli Reshotho and
Rudolph C. Haefeli. October 1952 35p aiagrs.,
photos., tab. (NACA RM E52H28)

The flow characteristics of three axially symmetric
multinozzles of nominal design Mach numbers 3.46,
3.07, and 7.01 and a two-dimensional multinozzle of
nominal design Mach number 3.07 were investigated.
Three types of disturbance, classified as oblique
shock waves, corner shock waves, and wakes, were
observed in the flow fields. The effect of the geom-
etry of the multinozzles on these disturbances and on
the actual Mach number and pressure recovery ob-
tained with the multinozzle is discussed herein. The
magnitude of the disturbances was found to depend
on the exit turning angle of the Individual nozzles.
Generally, Mach number variations of 4-1. 2 percent
or less existed in a region which could be used as
a test section. The use of multinozzles, however,
appears to be restricted to experiments in which
nonuniformity of flow and large pressure losses can
be tolerated and for which simple fabrication and
quick Interchange of nozzles are desired.


NACA RM E52J31

EXPERIMENTAL INVESTIGATION OF RADOME
ICING AND ICING PROTECTION. James P. Lewis
and Robert J. Blade. January 1953. 60p. diagrs.,
photos. (NACA RM E52J31)

In an investigation of radome icing and icing protec-
tion in the NACA Lewis icing research tunnel, the
impingement of water and the formation of ice on two
radome configurations were found to agree well with
theory and experience. The ice formations on the
radomes produced serious effects on radar perform-
ance. The ethylene glycol fluid-protection system
gave adequate Icing protection for both anti-icing and
de-icing. The radomes were investigated at air-
speeds up to 290 miles per hour, air total tempera-
tures of -150 to 200 F, water contents up to 1.0
gram per cubic meter, and angles of attack of 00 and
40.








20

NACA RM E52L15a

PRELIMINARY INVESTIGATION OF ZIRCONIUM
BORIDE CERAMALS FOR GAS-TURBINE BLADE
APPLICATIONS. Charles A. Hoffman April 1953.
13p. photos., diagrs.. 2 tabs. (NACA RM E52L15a)


Zirconium boride ZrB2 ceramals were investigated
for possible gas-turbine-blade application. Included
in the study were thermal shock evaluations of disks,
preliminary turbine-blade operation, and observa-
tions of oxidation resistance. Thermal shock disks
of the following three compositions were studied:
(a) 97. 5 percent ZrB2 plus 2. 5 percent B by weight;
(b) 92. 5 percent ZrB2 plus 7. 5 percent B by weight;
and (c) 100 percent ZrB2. Thermal shock disks
were quenched from temperatures of 18000, 20000,
22000, and 24000 F. The life of turbine blades con-
taining 93 percent ZrB2 plus 7 percent B by weight
was determined in gas-turbine tests. The blades
were run at approximately 16000 F and 15, 000 to
26, 000 rpm. The thermal shock resistance of the
97.5 percent ZrB2 plus 2. 5 percent boron ceramals
compares favorably with that of TiC plus Co and TIC
plus Ni ceramals. Oxidation of the disks during the
thermal shock evaluation was slight for the compara-
tively short time (8. 3 hr) up through 20000 F. Oxi-
dation of a specimen was severe, however, after
100 hours at 20000 F. The turbine blade perform-
ance evaluation of the 93 percent ZrB2 plus 7 percent
B composition was preliminary in scope; no con-
clusions can be drawn.


NACA RM E52L19a

EXPERIMENTAL INVESTIGATION OF A LIGHT-
WEIGHT ROCKET CHAMBER. John E. Dalgleish
and Adelbert 0. Tischler. March 1953. 12p.
photos. (NACA RM E52Ll9a)

Experiments have been conducted with a jacketed
rocket combustion chamber that was fabricated by
hydraulic-forming from sheet metal Rocket com-
bustion chambers made by this method have been
used successfully. Runs with these combustion
chambers have been made at over-all heat-transfer
rates of 1. 7 Btu per square inch per second with
water cooling and also with ammonia as a regenera-
tive coolant.


NACA RM E52L31

THERMAL-SHOCK RESISTANCE OF A CERAMIC
COMPRISING 60 PERCENT BORON CARBIDE AND
40 PERCENT TITANIUM DIBORIDE. C. M.
Yeomans and C A. Hoffman. March 1953. 7p.
photos.. 3 tab. (NACA RM E52L31)

Thermal-shock resistance of a ceramic comprising
60 percent boron carbide and 40 percent titanium
diboride was investigated. The material has ther-
mal shock resistance comparable to that of NBS body
4811 IC and that of zirconia, but is inferior to
beryllia., alumina, and titanium-carbide ceramals.
It is not considered suitable for turbine blades.


NACA
RESEARCH ABSTRACTS NO. 56

NACA RM E53A21

FUEL CHARACTERISTICS PERTINENT TOpTE
DESIGN OF AIRCRAFT FUEL SYSTEMS. Henry-C.
Barnett and R. R. Hibbard. June 1953. 1, 104p. '
diagrs., 9 tabs. (NACA RM E53A21)

Because of the importance of fuel properties in
design of aircraft fuel systems the present report
has been prepared to provide information on the
characteristics of current Jet fuels. In addition to
information on fuel properties, discussions are
presented on fuel specifications, the variations
among fuels supplied under a given specification,
fuel composition, and the pertinence of fuel composi-
Lion and physical properties to fuel system design.
In some Instances the Influence of variables such as
pressure and temperature on physical properties is
indicated. References are cited to provide fuel
system designers with sources of Information
containing more detail than is practicable In the
present report.


NACA RM E53E25

ANALYTICAL INVESTIGATION OF FUEL TEMPER-
ATURES AND FUEL-EVAPORATION LOSSES EN-
COUNTERED IN LONG-RANGE HIGH-ALTITUDE
SUPERSONIC FLIGHT. Richard J. McCafferty.
August 1953. 38p. diagrs., 2 tabs. (NACA
RM E53E25)

Two types of fuel losses were considered, loss due to
adiabatic evaporation during climb and loss due to
aerodynamic-heating effects. Heat-balance relations
were employed to estimate amount of heat trans-
ferred to fuel contained in a cylindrical fuselage.
Influence of flight speed, flight altitude, fuel with-
drawal rate, tank size, tank pressurization, insula-
tion, and initial fuel temperature on evaporation of
JP-4 fuel was predicted. Influence of fuel volatility
was determined by comparing evaporation obtained
with JP-4 fuel and that obtained with JP-5. Results
of analysis predicted, for an uninsulated tank, JP-4
fuel-boiling losses of 38 percent of total fuel weight
carried by aircraft at the maximum flight speed and
altitude considered. Fuel losses increased with in-
creasing flight speed and flight time but decreased
with increasing altitude. Fuel loss caused by adiabat-
ic evaporation could be eliminated by ground refrig-
eration, tank pressurization, or use of low volatility
fuel. Fuel loss due to aerodynamic heating could be
eliminated by a combination of insulation of a prac-
tical thickness and use of a low volatility fuel; how-
ever, if tank pressurization were employed without
insulation excessive pressures would be required.
Without Insulation, a maximum liquid-fuel tempera-
ture of 7400 F would result if no fuel cooling oc-
curred as a result of fuel evaporation; high liquid-
fuel temperatures may create additional fuel handling
problems.

NACA RM E53FO3

THE SYNTHESIS OF BUTYLSILANES BY A LARGE-
SCALE REDUCTION WITH LITHIUM ALUMINUM
HYDRIDE. Samuel Kaye, Stanley Tannenbaum and
Harold F. Hipsher. July 1953. 12p. photos.,
diagr. (NACA RM E53F03)







NACA
RESEARCH ABSTRACTS NO. 56

Eighty pounds of butylsilanes required for test pur-
poses was synthesized by reducing butyltrichlorosil-
ane with Lithium aluminum hydride in diozane solu-
tion. This was safely and successfully accom-
plished by operating in steel equipment under an at-
mosphere of oil -pumped nitrogen to minimize the
hazards involved in conducting this operation on a
large scale. The apparatus and procedure used to
carry out the necessary manipulations are described
in this report.

NACA RM E53F26

THERMAL SHOCK RESISTANCE AND HIGH-
TEMPERATURE STRENGTH OF A MOLYBDENUM
DMILICIDE ALUMINUM OXIDE CERAMIC. W. A.
Maxwell and R. W. Smith. October 1953. 8p.
photo., 4 tabs. (NACA RM E53F26)

A ceramic having the nominal composition 75 percent
molybdenum disilicide, 25 percent aluminum oxide
was evaluated in thermal shock by two methods and
high-temperature modulus-of -rnipture strengths were
determined. The material was found to have im-
proved thermal-shock resistance as compared with
pure molybdenum disilicide and to have moderate
high-temperature strength.

NACA RM E53I16

FUEL CHARACTERISTICS PERTINENT TO THE
DESIGN OF AIRCRAFT FUEL SYSTEMS. SUPPLE-
MENT I ADDITIONAL INFORMATION ON
MIL-F-7914(AER) GRADE JP-5 FUEL AND
SEVERAL FUEL OILS. Henry C. Barnett and
Robert R. Hibbard. (Supplement to RM E53A21)
November 1953. 39p. diagre., 3 tabs. (NACA
RM E53116)

Since the release of the first NACA publication of
fuel characteristics pertinent to the design of air-
craft fuel systems, additional information has be-
come available on MIL-F-7914(AER) grade JP-5
fuel and several of the current grades of fuel oils.
In order to make this information available to fuel-
system designers as quickly as possible, the present
report has been prepared as a supplement to NACA
RM E53A21, "Fuel Characteristics Pertinent to the
Design of Aircraft Fuel Systems' by Henry C.
Barnett and R. R. Hibbard, 1953.


NACA RM L50A04a

MAXIMUM-LIFT CHARACTERISTICS OF A WING
WITH THE LEADING-EDGE SWEEPBACK DE-
CREASING FROM 450 AT THE ROOT TO 200 AT
THE TIP AT REYNOLDS NUMBERS FROM 2.4 x 10
TO 6.0 a 106. Roy H. Lange. July 6, 1950. 62p.
diagrs., photos. (NACA RM L50A04a)


Results are presented of an investigation of the max-
imum lift characteristics of a wing with the leading-
edge sweepback decreasing from 450 at the root to
20 at the tip. The investigation was made for con-
ditions of leading edge smooth and leading edge
rough for the basic wing and for the wing with split
flaps, leading-edge flaps, outboard slats, and combi-
nations of these hi@-lift devices at Reynolds num-
bers from 2.4 x 100 to 6.0 x 106.


21

NACA RM LSOA23a

LA -MmPEED PRESSURE-DSTRIBUTrON MEASURE-
MENTS AT A REYNOLDS NUMBER OF 3.5 x 106 ON
A WING WITH LEADING-EDGE SWEEPBACK DE-
CREASING FROM 450 AT THE ROOT TO 2P AT
THE TIP. U. Reed Barnett, Jr. and Roy H. Lange.
July 7, 1950. 39p. diagre., photo. (NACA
RM L50A23a)


Results are presented of an investigation to deter-
mine the pressure distributions over a wing with the
leading-edge sweepback decreasing from 450 at the
root to 20P at the tip. Tests were made at a
Reynolds number of 3.5 z 106 and a Mach number of
0.07 on the wing with and without 60 split flaps.


NACA RM L50CO2

LOW-SPEED PITCHING DERIVATIVES OF LOW-
ASPECT-RATIO WINGS OF TRIANGULAR AND
MODIFIED TRIANGULAR PLAN FORMS. Alex
Goodman and Byron M. Jaquet. April 17, 1950. 25p.
diagram photos., 2 tabs. (NACA RM L50C02)


The present paper presents results of an investiga-
tion made in the Langley stability tunnel to deter-
mine effects of change in profile and aspect ratio on
pitching derivatives of triangular and modified trian-
gular wings. Results were compared with available
theories.

NACA RM L50C15a

EXPERIMENTAL INVESTIGATION OF THE EFFECT
OF ASPECT RATIO AND MACH NUMBER ON THE
FLUTTER OF CANTILEVER WINGS. E. Widmayer,
Jr., W. T. Lauten, Jr. and S. A. Clevenson.
June 1, 1950. 20p. diagrs., 2 tabs. (NACA
RM L50C15a)


The results of some wind-tunnel experiments to in-
vestigate the effects of aspect ratio and Mach number
on the flutter of uniform, unswept, cantilever wings
are reported. Models having aspect ratios ranging
from 2 to 13 were tested at Mach numbers up to 0.92.
No general attempt is made to correlate the data with
three-dimensional-flow theory, but an examination of
the data is made on the basis of reference theoretical
values obtained from the two-dimensional incompres-
sible flow theory. On this basis a reduction in
aspect ratio, in general, increased the ratio of the
experimental flutter speed to the calculated flutter
speed. The analysis also indicated that for a given
aspect ratio, this ratio decreased slightly as the
Mach number is increased.



NACA RM L50D06

ABILITY OF PILOTS TO CONTROL SIMULATED
SHORT-PERIOD YAWING OSCILLATIONS. William
H. Phillips and Donald C. Cheatham. November 13,
1950. 23p. diagrs., photos., tab. (NACA
RM 150D06)







22

The results of an Investigation of pilot's ability to
control short-period yawing oscillations in a simulat-
ing device are presented. Pilot's ability to control
the short-period yawing oscillations has been deter-
mined as a function of period, control effectiveness,
and inherent damping or instability.




NACA RM L50D07

PRESSURE-RISE AND LEAKAGE-LOSS CHARAC-
TERISTICS OF A ROTATING COWLING. Jack F.
Runckel and Gerald Hieser. August 30, 1950. 47p.
diagrs., photos. INACA RM L50D07)


Presents results of an investigation of pressure-rise
and leakage-loss characteristics of a rotating cowl-
ing. A method for predicting the pressure rise is
developed and compared with the results obtained
experimentally. Design considerations peculiar to
this type of cowling are discussed, and methods for
estimating the leakage flow and proper setting of the
cowling vane are outlined.



NACA RM L50ID13

A WIND-TUNNEL INVESTIGATION OF THE EF-
FECTS OF THRUST-AXIS INCLINATION ON PRO-
PELLER FIRST-ORDER VIBRATION. W. H. Gray,
J. M. Hallissy, Jr. and A. R. Heath, Jr. June 9,
1950. 64p. diagrs., photo., tab. (NACA RM L50DIS)


Data are presented on aerodynamic excitation of pro-
peller vibration with the thrust axis at two angles of
Inclination to the air stream. Aerodynamic excita-
tion is calculated and compared with measured values
for several conditions. Measured and calculated
blade vibratory stresses for several excitation values
are presented. The calculated values of exciting
force compared well with the measured values at low
rotational speeds, not so well at high rotational
speeds. First-order blade vibratory stresses were
computed with satisfactory accuracy from unpitched
propeller loading data.




NACA RM L50D14

LOW-SPEED LATERAL STABILITY AND AILERON-
EFFECTIVENESS CHARACTERISTICS AT A
REYNOLDS NUMBER OF 3.5 x 106 OF A WING WITH
LEADING-EDGE SWEEPBACK DECREASING FROM
450 AT THE ROOT TO 200 AT THE TIP. Roy H
Lange and Huel C. McLemore. July 6, 1950. 44p.
dzagrs.. photos. (NACA RM L50D14)


Results are presented of an investigation of the static
lateral stability and aileron-effectiveness character-
istics of a wing with the leading-edge sweepback de-
creasing from 450 at the root to 20 at the tip. The
investigation is made for the basic wing and for the
wing with split flaps, leading-edge flaps, outboard
slats, and combinations of these high-lift devices at
a Reynolds number of about 3.5 x 106.


NACA
RESEARCH ABSTRACTS NO. 56

NACA RM L50E02

LOW-SPEED INVESTIGATION OF LEADING-EDGE
AND TRAILING-EDGE FLAPS ON A 47.50 SWEPT-
BACK WING OF ASPECT RATIO 3.4 AT A .
REYNOLDS NUMBER OF 4.4 x 106. Jerome
Pasamanick and Thomas B. Sellers. June 12. 1950.
29p. diagrs.. Dhoto. (NACA RM L50E02)


Presents results of an investigation In the Langley
full-scale tunnel of various extensible leading-edge
and plain trailing-edge flaps on a wing-fuselage com-
bination. The wing leading-edge sweep was 47.50,
the aspect ratio was 3.4, the taper ratio was 0.51, and
the airfoil sections were NACA 641A112. The data
include the effects of flap design parameters on the
lift, drag, and pitching-moment characteristics at
zero yaw for a range of angle of attack at a Reynolds
number of 4.4 x 100.


NACA RM L50EI2a

EFFECT OF VARIOUS OUTBOARD AND CENTRAL
FINS ON LOW-SPEED YAWING STABILITY
DERIVATIVES OF A 600 DELTA-WING MODEL.
Alex Goodman. June 19, 1950. 35p. diagrs.,
photos., 2 tabs. (NACA RM L50E12a)


The results of an investigation conducted in the
Langley stability tunnel to determine the effects of
various outboard- and central-fin arrangements on
the low-speed yawing stability derivatives of a 600
delta-wing model are presented. Calculated values
of the effective center of pressure for several of the
central fins are also presented. Procedures are
suggested for estimating the contribution of fins to
the stability derivatives of similar model configura-
tions.

NACA RM L50Fi6

LOW-SPEED AERODYNAMIC CHARACTERISTICS
OF A SERIES OF SWEPT WINGS HAVING NACA
65A006 AIRFOIL SECTIONS. (Revised) Jones F.
Cahill and Stanley M. Gottlieb. October 17, 1950.
63p. diagrs., photos. (NACA RM L50FI6. Formerly
RM L9J20)


An investigation was made to determine the effect of
sweep, taper ratio, and aspect ratio on the aero-
dynamic characteristics of nine semispan wings of
NACA 65A006 airfoil section with and without split
flaps. Lift, drag, pitching-moment, and wing-root
bending-moment characteristics were measured
through a range of Reynolds numbers from 1.5 x 10"
to 12.0 x 106. One of these wings was tested with a
hinged leading-edge flap of various spans and deflec-
tions to determine the effect of this type of flap on
longitudinal stability near maximum lift.

NACA RM L50F16a

LOW-SPEED LONGITUDINAL CHARACTERISTICS
OF A CIRCULAR-ARC 520 SWEPTBACK WING OF
ASPECT RATIO 2.84 WITH AND WITHOUT
LEADING-EDGE AND TRAILING-EDGE FLAPS AT
REYNOLDS NUMBERS FROM 1.6 x 106 TO 9.7 x 106.
Gerald V. Foster and Roland F. Griner. August 11,
1950. 40p. diagrs., photo., tab. (NACA RM L50F16a)







NACA
RESEARCH ABSTRACTS NO. 56

The results of low-speed tests of a 52o swept wing
which had circular-arc sections are presented. The
wing exhibited Large changes in stability from the low
to high lift range. The stability of the wing was im-
proved with the addition of leading-edge flaps which
extended over 25 percent of the wing semispan. The
aerodynamic characteristics of the wing were not
appreciably affected by the variation of Reynolds
numbers.


NACA RM L50G17

WIND-TUNNEL INVESTIGATION OF THE EDW-
SPEED LONGIUDINAL AND LATERAL CONTROL
CHARACTERISTICS OF A TRIANGULAR-WING
MODEL OF ASPECT RATIO 2.31 HAVING
CONSTANT-CHORD CONTROL SURFACES. Walter
plWomlhart and9 Wlliam H. Michael,pho/.,2ab.
(NACA RM L50G17)


Results are presented of a wind-tunnel investigation
of lowr-speed longitudinal and lateral control charac-
teristics of a model with a triangular wing having
NACA 65(06)-006.5 airfoil sections and aspect ratio
2.31 and equipped with constant-chord control enr-
faces. The investigation included determination of
effects of adding fuselage to wing, adding transition
strips, varying the Reynolds number from 1.62 x 106
to 2.62 x 1 ,c1 and sealing the gap at the nose of the



NACA RMg L50B2

POSITIONIN INVESTIGATION OF SINGLE
SIUrrTED FLAPS ON A 4f.70" SWEPTBACK WING
AT REY LD~lS NUMBERS OF 4.0 x 106 AND
6.0 x 10 Stanley H. Spooner and Ernet F.
Mollenberg. October 9, 1950. 36p. diagra., photo.'
3 tabs. (NACA RM L50H29)


The relationship between flap effectiveness and the



effects of Reynolds number variation and a compari-
son with two-dimensional posilloning studies of a
slotted flap. The low-speed investigation was con-
ducted in the Langley 10-foot pressure tunnel at
Reynolds numbers of 4.0 x 106 and 6.0 x 106-


NACA RM L50I29


DEWICLET MERMENTLON FLOAERR INOT1VING
WING DEFORMATION AND BODY MOTIONS. H. J.
Cunningham and R. R. Lundstrom. November 30,
1950. 27p. diagrs., photos., 2 tabs. (NACA
RM L50I20)


Flight tests and a mathematical analysis were car-
rted out to demonstrate and confirm a type of sub-
sonic flutter involving rigid-body motions and wing
deformations. For the configuration considered, the
period of the oscillation was approximately 100
chords per cycle which la well within the range of


23

period found in dynamic-staballty work on rigid air-
craft with free controls. A mathematical analyeze
based on two-dimensional incompressible flow prol-
vided a conservative prediction of the airspeed at
which the low-frequency flutter occurred. RI was
found that wing-bending stiffness is the Important
parameter for preventing such flutter.



NACA RM L50L11

THEORETICAL AND ANALOG STUDIES OF THE
EFFECTS OF NONLINEAR STABILITY DERIVA-
TIVES ON THE LONGITUDINAL MOTIONS OF AN
AIRCRAFT IN RESPONSE TO STEP CONTROL DE-
FLECTIONS AND TO THE INFLUENCE OF PRO-
PORTIONAL AUTOMATIC CONTROL. HowarrdJ.
Curtfman, Jr.CAFe ruar023, 1951. 55p.diagre.,


Through theoretical and analog results the effects of
two nonlinear stability derivatives on the longitudnal
motions of an aircraft have been investigated. Non-
linear functions of pitching-moment and 111 coeffi-
cients with angle of attack were considered. Analog
results of aircraft motions in response to step ele-
vator deflections and to the action of the proportional
control systems are presented. The occurrence of
continuous hunting oscillations was predicted and
demonstrated for the attitude stabilization system
with proportional control for certain nonlinear
pitching-moment variations and autopilot adjust-
ments.


NACA RM L50Ll2b

THE EFFECTS OF VARIOUS PARAMETERS IN-
CLUDING MACH NUMBER ON PROPELLER-BLADE
FLUTTER WITH EMPHASIS ON STALL FLUTTER.
John E. Baker. January 31, 1951. 4l0p. diagrs.,
3 tabs. (NACA RM L50Ll2b)





respect to propeller stall flutter. The minimum
values of the flutter-speed coefficient were found to
be shghtly greater than 1.0 at subcritical Mach num-
bers. Of the fewr parameters that raised the mim-
mum flutter-speed coefficients, forward movement
of the section center-of-gravity location and Mach
number at supercritical speeds were most eiguft-
cant. The effect of Mach number was of such sig-
nificance that a tentative criterion for designing
completely flutter-free thin supersonic propellersr




NACA RM L51Al0

AN INVESTIGATION OF THE EFFECT OF
VERTICAL-FIN LOCATION AND AREA ON LOW-
SPEED LATERAL STABILITY DERIVATIVES OF A
SEMITAILLESS AIRPLANE MODEL. LewrisR.
Fisher and William H. Michael, Jr. March 7, 1951.
41p. diagram photos. (NACA RM LStA10)








24
A low-speed wind-tunnel investigation was conducted
in order to determine the effect of vertical-fin loca-
tion and area on the static and rotary lateral sta-
bility derivatives of a semitailless airplane model
and to establish the validity of the application of the
usual methods for predicting the stability derivatives
of such airplanes. Comparisons are made of the
experimental and calculated derivatives for the wing
and vertical fins wherein consideration is taken of
some interference effects.
NACA RM L51A26

INVESTIGATION AT LOW SPEED OF THE EFFEC-
TIVENESS AND HINGE MOMENTS OF A CONSTANT-
CHORD AILAVATOR ON A LARGE-SCALE TRIAN-
GULAR WING WITH SECTION MODIFICATION.
John G. Hawes and Ralph W. May, Jr. April 24,
1951. 47p. diagrs., photos., tab. (NACA
RM L51A26)


Results are presented of an investigation to deter-
mine the low-speed longitudinal, lateral, and hinge-
moment control characteristics at zero yaw of a
large-scale 600 delta wing having a 10-percent -thick
biconvex symmetrical airfoil. The wing was also
tested with a nose glove having NACA 65-010 section
ordinates. The wing was equipped with a constant-
chord plain-type semispan ailavator having two
segments.

NACA RM L51B09

WIND-TUNNEL INVESTIGATION AT LOW SPEED
OF THE EFFECTS OF SYMMETRICAL DEFLEC-
TION OF HALF-DELTA TIP CONTROLS ON THE
DAMPING IN ROLL AND YAWING MOMENT DUE TO
ROLLING OF A TRIANGULAR-WING MODEL.
Walter D. Wolhart. April 6, 1951. 17p. diagrs.,
photos. (NACA RM L51B09)


Results are presented of a wind-tunnel investigation
at low speed to determine the effects of symmetrical
deflection of half-delta tip controls on the damping in
roll and yawing moment due to rolling of a triangular
wing model. The investigation included determina-
tion of the effects of control size as well as control
deflection.


NACA RM L51D20

LOW-SPEED INVESTIGATION OF SEVERAL TYPES
OF SPLIT FLAP ON A 47.70 SWEPTBACK-WING -
FUSELAGE COMBINATION OF ASPECT RATIO 5.1
AT A REYNOLDS NUMBER OF 6.0 x 106. Stanley
H. Spooner and Ernst F. Mollenberg. July 1951.
41p. diagrs., photo., tab. (NACA RM L51D20)


Results are presented of an investigation to deter-
mine the characteristics of a 47.70 swept back-wing -
fuselage combination equipped with split flaps and
several modifications thereof. The modifications
consisted of extended split (Zap), rotated split, step
split, and triangular flaps. The wing had an aspect
ratio of 5.1, a taper ratio of 0.383, and NACA 64-210
airfoil sections. The maximum wing thickness in a
plane parallel to the plane of symmetry was 7.5 per-
cent of the wing chord. The tests were conducted in
the Langley 19-foot pressure tunnel at a


NACA
RESEARCH ABSTRACTS NO. 56
Reynolds number of 6.0 x 106 and a Machi number
of 0.14.


NACA RM L51D20a .
.
LOW-SPEED STATIC LONGITUDINAL STABILITY
AND CONTROL CHARACTERISTICS OF A 600
TRIANGULAR-WING MODEL HAVING HALF-
DELTA TIP CONTROLS. Byron M. Jaquet, M. J.
Queijo and Jacob H. Lichtenstein. June 1951. 30p.
diagrs., photos. (NACA RM L51D20a)


The results are presented of an investigation to de-
termine the low-speed static longitudinal stability
and control-effectiveness characteristics ofat a 600
triangular-wing model having half-delta tip controls.
Comparisons of results are made with constant-
chord flaps. Effects of fuselage and end plates on
stability and control characteristics are presented.
Results are compared with available theory where
possible.


NACA RM L51EIla

EFFECTS OF INLET WALL CONTOUR ON THE
PRESSURE RECOVERY OF A 100 10-INCh-INLET-
DIAMETER CONICAL DIFFUSER. Martin R. Copp.
September 1951. 29p. diagrs., photos. (NACA
RM L51Ella)

An investigation was made with a 100 10-inch-inlet-
diameter conical diffuser of 2:1 area ratio to deter-
mine the influence on flow characteristics and pres-
sure recovery of sharpness of curvature of the junc-
tion between the walls of a conical diffuser and a
cylindrical approach tube. The investigation was
performed for two thicknesses of inlet boundary
layer over a Reynolds number range of 1.2 x 106 to
3.2 x 106 and a mean inlet Mach number range of
approximately 0.28 to choking. The over-all static-
pressure recovery, expressed as the ratio of the
actual rise in static pressure divided by the ideal
rise in static pressure, was entirely
independent of the inlet wall shape.


NACA RM L51E16

WING-FLOW INVESTIGATION OF A 450 CONE AS
AN ANGLE-OF-ATTACK MEASURING DEVICE AT
TRANSONIC SPEEDS. Herbert C. McClanahan, Jr.
July 1951. 15p. photos., diagrs. (NACA RM
L51E16)

A study of pressure measurements on a cone of 450
apex angle made at angles of attack from -100 to 500
and zero yaw angle through a Mach number range of
0.7 to 1.1 to determine the usefulness of a cone for
measuring angle of attack at transonic speeds.

NACA RM L51F22

INVESTIGATION OF LOW-SPEED LATERAL
CONTROL AND HINGE-MOMENT CHARACTER-
ISTICS OF A 20-PERCENT-CHORD PLAIN
AILERON ON A 47.70 SWEPTBACK WING OF
ASPECT RATIO 5.1 AT A REYNOLDS NUMBER OF
6.0 x 106. William M. Hadaway and Reino J. Salmi.
October 1951. 31p. diagrs. (NACA RM L51F22)







NACA
RESEARCH ABSTRACTS NO.56
Contains results of wind-tunnel tests to determine
the lateral control, hinge-moment, aileron-load, and
balance-chamber-pressure characteristics of a 20-
percent-chord plain aileron on a 47.70 sweptback
wing of aspect ratio 5.1 having NACA 64-210 airfoil
sections with and without high-lift and stall-control
devices. The investigation was conducted at a
Reynolds number of 6.0 x 106 and a corresponding
Mach number of 0.14 for an angle-of-attack range of
-40 to approximately 300.

NACA RM L51G30

A COMPARISON OF THE SPANWISE LOADING CAL-
CULATED BY VARIOUS METHODS WITH EXPERI-
MENTAL LOADINGS OBTAINED ON A 450 SWEPT-
BACK WING OF ASPECT RATIO 8 AT A REYNOLDS
NUMBER OF 4. 0 x 106. William C. Schneider.
January 1952. 32p. diagrs., tab. (NACA RM
L51G30)

Experimental force and moment data obtained by
pressure measurements on a wing of aspect ratio
8.02, 450 sweepback of the quarter-chord line, taper
ratio of 0.45, and NACA 63 lA01l2 airfoil sections
parallel to the plane of symmetry have been com-
pared with the calculated loadings obtained by the
standard methods proposed by Weissinger, Falkner,
and Multhopp, as well as by variations thereof On
the basis of the present calculations, the Weissinger
solution, when the number of control points was in-
creased from 7 (the number suggested by Weissinger)
to 15, or the Multhopp solution when using at least
15 control points, resulted in a good compromise
between lift-curve slope and load shape, although all
of the calculations underestimated CL .


NACA RM L51HO9a

EXPERIMENTAL AND ANALYTICAL INVESTIGA-
TION OF FLUTTER OF A NONUNIFORM SWEPT-
BACK CANTILEVER WING WITH TWO CONCEN-
TRATED WEIGHTS. John L. Sewall. December
1951. 33p. diagrs., photos., 3 tabs. (NACA RM
L51H09a)

An experimental and analytical investigation of the
flutter of a weighted cantilever wing, swept back
34.50, is reported. Concentrated weights which were
78.4 percent and 42.0 percent of the weight of the
wing were mounted at approximately one-third and
three-quarters of the span respectively, and flutter
data obtained for this weighted condition were ana-
lyzed by means of a Rayleigh-Ritz type of flutter
theory employing uncoupled still-air vibration modes.
Unusually good agreement was obtained by using two
modes (fundamental bending and fundamental torsion),
and the introduction of three and four higher modes
did not appreciably affect this agreement. Additional
information related to the theory is given in an
appendix.

NACA RM L51HI5

COMPARISON OF SEMISPAN AND FULL-SPAN
TESTS OF A 47.50 SWEPTBACK WING WITH SYM-
METRICAL CIRCULAR-ARC SECTIONS AND HAV-
ING DROOPED-NOSE FLAPS, TRAILING-EDGE
FLAPS, AND AILERONS. Stanley Lipson and U. Reed
Barnett, Jr. December 1951. 60p. photos., diagrs.
(NACA RM L51H15)


25
An investigation was conducted in the Langley full-
scale tunnel to determine the comparative charac-
teristics of a full-span and semispan 47.50 sweptback
wing having 10-percent thick biconvex airfoil sections
and an aspect ratio of 3.5. The configurations tested
were the basic wing, the wing with various combina-
tions of leading- and trailing-edge flaps, and the
basic wing with an aileron deflected. The results In-
cluded both force and pressure measurements and
were obtained at a Reynolds number of 4.2 x 106 and
a Mach number of 0.07.

NACA RM L51H31

PRELIMINARY INVESTIGATION OF THE TRANSFER
OF HEAT FROM A FLAT PLATE AT A MACH NUM-
BER OF 1.5. M. A. Emmons, Jr. and R. F.
Blanchard. December 1951. 23p. diagrs., photos.
(NACA RM L51H31)


Surface temperatures and heat transfer to the air
stream have been measured for turbulent flow over a
flat plate at a Mach number of 1.5 and at a Reynolds
number, based on the momentum thicknesses of
boundary layer, of approximately 5000. Preliminary
data are presented and the surface heat-transfer co-
efficient calculated from these data is considered to
be accurate to t2.6 percent at a temperature poten-
tial of 500 F. These data are in good agreement with
the results produced by applying modificati ins ob-
tained from published information to existing sub-
sonic theories.


NACA RM L51110

LOW-SPEED WIND-TUNNEL INVESTIGATION OF
LATERAL CONTROL CHARACTERISTICS OF A 600
TRIANGULAR-WING MODEL HAVING HALF-DELTA
TIP CONTROLS. Byron M. Jaquet and M. J. Queijo.
November 1951. 50p. diagrs., photos. (NACA RM
L51110)

Results are presented of a low-speed investigation of
the lateral control'characteristics of a 600 triangular
wing model equipped with half -delta tip controls hav-
ing areas of 5, 10, and 15 percent of the wing area
(sum of left- and right-control areas). Effects of
end plates are determined for 10-percent-area con-
trols. Half -delta tip controls are compared with
constant-chord controls. Experimental control and
rolling effectiveness are compared with theory.

NACA RM L51J05

LOW-SPEED INVESTIGATION OF THE EFFECTS
OF WING LEADING-EDGE MODIFICATIONS AND
SEVERAL OUTBOARD FIN ARRANGEMENTS ON
THE STATIC STABILITY CHARACTERISTICS OF A
LARGE-SCALE TRIANGULAR WING. H. Clyde
McLemore. January 1952. 64p. diagrs., photo.,
tab. (NACA RM L51JO5)

Results are presented of an Investigation In the
Langley full-scale tunnel to determine the effects of
wing leading-edge modifications and several outboard
fin arrangements on the low-speed static-stability
characteristics of a large-scale triangular wing
having 600 of leading-edge sweep and 10-percent-
thick circular-arc airfoil sections parallel to the
plane of symmetry. The R olds number range was
from 2.70 x 106 to 9.72 x 10 .







26

NACA RM L51J19

THEORETICAL CALCULATION OF THE EFFECT OF
THE FUSELAGE ON THE SPANWISE LIFT DISTRI-
BUTION ON A WING. Martin Ziotnick and Franklin
W. Diederich. March 1952. 27p. diagrs. (NACA
RM L51J19)

A method is presented for calculating the effects of
the fuselage on the spanwise lift distribution on a
wing by an application of a conformal-mapping proce-
dure to the simplified lifting-surface theory. This
method is applicable to any symmetrical wing-
fuselage configuration. An outline of the computing
procedure and a comparison of the calculated results
with experimental results are included. A numerical
example is given.

NACA RM L51J23

METHOD OF ESTIMATING THE STICK-FIXED
LONGITUDINAL STABILITY OF WING-FUSELAGE
CONFIGURATIONS HAVING UNSWEPT OR SWEPT
WINGS. Milton D. McLaughlin. January 1952. 41p.
diagrs., 3 tabs. (NACA RM L51J23)


A method is given for calculating the stick-fixed
longitudinal stability of a wing-fuselage configuration
at suberitical Mach numbers. The method applies to
unswept- and swept-wing configurations. A variation
of the additional fuselage loading due to the presence
of the wing is presented. A comparison of the calcu-
lated neutral points with neutral-points obtained from
experiment for 23 wing-fuselage configurations is
presented.

NACA RM L51J25

ANALYSIS OF V-g RECORDS FROM THE GRUMMAN
F8F-2 AIRPLANE. James 0. Thornton. July 1952.
lIp. diagrs., tab. (NACA RM L51J25)

V-g records obtained in training and operational
flights of the F8F-2 airplane are analyzed statisti-
cally to show the frequency with which large values
of normal acceleration and airspeed are encountered.
A comparison is made with previous V-g results
from the F8F-1 airplane.

NACA RM L51K09

THE EFFECT OF SURFACE ROUGHNESS ON THE
PERFORMANCE OF A 230 CONICAL DIFFUSER AT
SUBSONIC MACH NUMBERS. Jerome Persh.
January 1952. 42p. diagrs., photos. (NACA RM
L51K09)

An investigation was conducted to determine the ef-
fect of surface roughness on the performance of a 2:1
area ratio, 230 conical diffuser with a constant-area
tail pipe about 3-1 '2 inlet diameters in length. The
inlet-boundary-layer thickness was of the order of
5 percent of the inlet diameter. The air flows used
in this investigation cover an inlet Mach number
range from 0.10 to 0.64 corresponding to Reynolds
numbers of 106 to 6 x 106 based on inlet diameter.
The surface of the diffuser was coated with cork
particles of a controlled size. Incremental bands of
roughness were removed from the downstream edge
after each series of pressure measurements were


NACA
RESEARCH ABSTRACTS NO. 56

made and the variation of diffuser performance with
percent of diffuser length roughened thereby deter-
mined.


- a'- *


NACA RM L51K26


FORCE AND PRESSURE INVESTIGATION AT
LARGE SCALE OF A 490 SWEPTBACK SEMISPAN
WING HAVING NACA 65A006 SECTIONS AND
EQUIPPED WITH VARIOUS SLAT ARRANGEMENTS.
Stanley Lipson and U. Reed Barnett, Jr. January
1952. 60p. diagrs., photo., tab. (NACA RM
L51K26)

The effects of slat span and slat deflection angle on
the longitudinal aerodynamic characteristics of a
semispan 49. 10 sweptback wing, aspect ratio 3. 78,
having NACA 65A006 airfoil sections have been inves-
tigated in the Langley full-scale tunnel. Normal-
force measurements were obtained on the extended
slat, both with and without a trailing-edge flap
deflected. The greater part of the data was obtained
at a Reynolds number of 6. 1 x 106 and a Mach num-
ber of 0. 10.

NACA RM L51L12

STUDIES OF THE FLOW FIELD BEHIND A LARGE
SCALE 47.50 SWEPTBACK WING HAVING
CIRCULAR-ARC AIRFOIL SECTIONS AND
EQUIPPED WITH DROOPED-NOSE AND PLAIN
FLAPS. Roy H. Lange and Marvin P. Fink. March
1952. 57p. diagrs.. photos., tab. (NACA RM
L51L12)

Results are presented of an investigation at large
scale of the downwash, sidewash, and wake charac-
teristics behind a 47 50 sweptback wing having
circular-arc airfoil sections and with drooped-nose
and plain flaps neutral and deflected 400. The sur-
veys were made at a Reynolds number of 4.3 x 106
and a Mach number of 0.07 and are presented for
three longitudinal distances behind the wing. Com-
parisons are made between the downwash calculated
by usine experimental span loadings and the experi-
mental downwash obtained.

NACA RM L52C17

A STUDY OF THE CHARACTERISTICS OF HUMAN-
PILOT CONTROL RESPONSE TO SIMULATED AIR-
CRAFT LATERAL MOTIONS. Donald C. Cheatham.
May 1952. 39p. diagrs., photos., tab. (NACA RM
L52C17)

There are presented studies of the characteristics of
pilot ability to control dynamically unstable yawing
oscillations, studies of pilot control response to
simulated aircraft yawing motions, and studies of the
feasibility of representing pilot control response in
an analytical form.

NACA RM L52C20

A PRELIMINARY GUST-TUNNEL INVESTIGATION
OF LEADING-EDGE SEPARATION ON SWEPT
WINGS. George L. Cahen. June 1952. 12p. photos.
(NACA RM L52C20)







NACA
RESEARCH ABSTRACTS NO.56


tufts on three wings having sweepback angles of 300,
450, and 600 show that under certain conditions a
leading-edge vortex can exist in the unsteady flow
associated with a gust It was indicated that, if a
wing in steady flight prior to entering a gust is at an
angle of attack several degrees less than that at which
vortex flow first begins in steady flow, it may pene-
trate the gust without having the vortex develop, even
if its angle of attack is increased by the gust into the
vortex-flow regime. However, if the wing is within
the vortex-flow regime prior to entering a gust, the
vortex flow can progress rapidly


NACA RM L52D04

LOW-SPEED WIND-TUNNEL INVESTIGATION OF
THE EFFECTS OF PROPELLER OPERATION AT
HIGH THRUST ON THE LONGITUDINAL STABILITY
AND TRIM OF A TWIN-ENGINE AIRPLANE CON-
FIGURATION. William C. Sleeman, Jr. and Edward
L. Linsley. July 1952. 66p. diagrs., photos.
(NACA RM L52D04)

An investigation was made to determine the effects of
dual-rotation propeller operation at high thrust on
the static longitudinal stability characteristics of a
semispan powered model representing a twin-engine
airplane configuration with flaps retracted. The
flow field behind the model was studied extensively
by several techniques. Dynamic pressure ratios and
effective downwash angles were obtained for a wide
range of tail heights and the general flow angularity
behind the model was obtained by the tuft-grid tech-
nique. Effects of extreme power on longitudinal
stability were evaluated for several combinations of
horizontal tall height and vertical location of the
center of gravity.


NACA RM L52DI7a

EFFECTS OF SEVERAL HIGH-LIFT AND STALL-
CONTROL DEVICES ON THE AERODYNAMIC
CHARACTERISTICS OF A SEMISPAN 490 SWEPT-
BACK WING U. Reed Barnett. Jr. and Stanley
Lipson. September 1952. 39p diagrs., photo
(NACA RM L52Dl7a|

Tests were conducted in the Langley full-scale tun-
nel of a wing with 49 10 sweepback of the leading
edge, aspect ratio 3.78, and NACA 65A006 sections
streamwise to determine the effect of a leading-edge
flap, a slat, and a plain trailing-edge flap on the
aerodynamic characteristics of a wing Limited
tests were also conducted concerning boundary-
layer control by suction and blowing. Rolling mo-
ments due to a flap-type aileron were also measured.
Most of the data were obtained at a Reynolds num-
ber of 6. 1 x 106 and a Mach number of 0 10

NACA RM L52D235a

EXPERIMENTAL INVESTIGATION OF FLOW
THROUGH THREE HIGHLY LOADED INLET GUIDE
VANES HAVING DIFFERENT SPANWISE CIRCULA-
TION GRADIENTS Loren A. Beatty, Melvyn
Savage and James C. Emery July 1952. 24p.
diagrs., lab. (NACA RM L52D25a)


27
Three highly loaded axial-flow compressor inlet
guide vanes with a free vortex and solid body design
were designed and tested in a low-speed annular cas-
cade A comparison net ween design and measured
turning angles was made by both neglecting and
taking into account the secondary flow effects Con-
tour plots of total pressure loss across the guide
vane were made. Wnen the radial gradient of circu-
lation was large (the hub axial velocity two or three
times greater than that at the tip), actual turning
angles were several degrees below design over the
outboard 30 percent of the vane. For such guide
vanes, the Lieblein and Ackley method for correct-
ing design two-dimensional turning angles by taking
into account secondary flows provided fairly good
agreement between design and measured turning
angles.



NACA RM L52H19

INVESTIGATION AT LOW SPEED OF THE DOWN-
WASH, SIDEWASH, AND WAKE CHARACTERISTICS
BEHIND A LARGE-SCALE TRIANGULAR WING,
INCLUDING THE EFFECTS OF YAW, FULL-SPAN
TRAILING-EDGE FLAPS, AND TWO LEADING-
EDGE MODIFICATIONS. Edward F. Whittle, Jr.
and John G. Hawes. October 1952. 65p. diagrs.,
photos., tab. (NACA RM L52H19)

Results are presented of a low-speed investigation of
the downwash, sidewash, and wake characteristics in
three planes behind a large-scale 600 triangular wing.
Included are the effects of yawing 100, full-span
plain flaps deflected 200, and two leading-edge nose-
glove modifications. Calculations of the static longi-
tudinal stability and trim characteristics of an
assumed 600 triangular-wing airplane configuration
showing the effect of three tail areas, two longi-
tudinal positions, and various vertical positions are
also presented.



NACA RM L52I24

SOME LOW-SPEED STUDIES OF THE EFFECTS OF
WING LOCATION ON WING-DEFORMATION-BODY-
FREEDOM FLUTTER. E. Widmayer, Jr.
November 1952. I21p. diagrs.,. 3 tabs. (NACA
RM L52124)

Some flutter experiments on unswept wings and 450
swept wings involving wing deformation and fuselage
mobility in pitch and performed at low air speeds
have been reported. The flutter was obtained for a
range of wing locations rearward of the pitching axis
and was characterized by low values of reduced fre-
quency. An analysis of the flutter has been made,
the flutter being calculated for several approxima-
tions to the aerodynamic forces. A comparison of
the results of the calculations with the results of the
experiments indicates that some accounting of the
effect of aspect ratio on the aerodynamic forces is
needed to adequately predict the flutter. The ex-
periments using the 450 swept wings yielded no
flutter for the range of wing locations and fuselage
inertias examined.







28

NACA RM L52J06

A CASCADE GENERAL-MOMENTUM THEORY OF
OPERATION OF A SUPERSONIC PROPELLER
ANNULUS. Bernard B. Klawans and Arthur W.
Vogeley. January 1953. 25p. diagram 2 tabs.
(NACA RM L52J06)

A cascade--general-momentum-theory method for
calculating the operating conditions of a supersonic
propeller annulus throughout the flight Mach number
range is presented. Representative subsonic, tran-
sonic, and supersonic solutions are given. The
method permits consideration of drag due to lift and
thickness, shock interference, and solidity and ap-
pears useful in studying the general trends of super-
sonic propeller operation.


NACA RM L52L30a

SUMMARY OF PILOTS' REPORTS OF CLEAR-AIR
TURBULENCE AT ALTITUDES ABOVE 10,000
FEET. Harry Press, Martin H. Schindler and
James K. Thompson. March 1953. 18p. diagrs.,
5 tabs. (NACA RM L52L30a)

A cooperative program for the collection of pilots'
reports of clear-air turbulence encountered at high
altitudes in normal civil and military operations was
undertaken in 1949. A simple postal-card question-
naire was distributed to groups of civil aircraft
operators and selected military units, and a sum-
mary of the data obtained from this survey is pre-
sented together with some additional reports from
Air Weather Service reconnaissance flights.

NACA RM L53A29

LOW-SPEED LATERAL CONTROL CHARACTERIS-
TICS OF AN UNSWEPT WING WITH HEXAGONAL
AIRFOIL SECTIONS AND ASPECT RATIO 4.0 AT A
REYNOLDS NUMBER OF 6.2 x 106. William M.
Hadaway. March 1953. 24p. diagrs., photo.
(NACA RM L53A29)

Results of wind-tunnel tests to determine the lateral
control,hinge moment, aileron load, and aileron
balance chamber pressure characteristics of both a
0. 79b '2 and 0.40b/2 outboard 25-percent-chord
aileron on a hexagonal 6-percent-thick wing of aspect
ratio 4. 0 and taper ratio 0.625 are presented. Tests
were made at a Reynolds number of 6. 2 x 106 and a
Mach number of 0.15 with a fuselage attached to the
wing. Effects of deflection of leading- and trailing-
edge flaps on the aileron characteristics of the out-
board aileron were also investigated.


NACA RM L53D09

LOW-SPEED, LARGE-SCALE INVESTIGATION OF
AERODYNAMIC CHARACTERISTICS OF A
SEMISPAN 490 SWEPTBACK WING WITH A
FOWLER FLAP IN COMBINATION WITH A PLAIN
FLAP, SLATS, AND FENCES Edward F. Whittle,
Jr. and Stanley Lipson June 1953. 42p. diagrs.,
photos., tab. (NACA RM L53D09)


NACA
RESEARCH ABSTRACTS NO.56

Tests were made in the Langley full-scale tunnel to
determine the effect of a Fowler flap on the aerody-
namic characteristics of a semispan 49.10 swept-
back wing having NACA 65A006 airfoil sections, an
aspect ratio of 3. 78, and a taper ratio of 0.59.
Slats, fences, and a plain flap were tested in corn-
bination with the Fowler flap, which was located
near the trailing edge of the plain flap. The tests
were conducted at Reynolds numbers of 6.1 x 106
and 4. 4 x 106 with corresponding Mach numbers of
0. 10 and 0. 07. respectively.



NACA RM L53E01a

MOMENT OF INERTIA AND DAMPING OF FLUID
IN TANKS UNDERGOING PITCHING OSCILLA-
TIONS. Edward Widmayer, Jr. and James R.
Reese. June 1953. Op. diagram. (NACA
RM L53E01a)

Studies were made of the fluid dynamics of some
different-shaped fuel tanks considered as being in a
horizontal flight attitude and undergoing pitching
oscillations of a few degrees. For the tank-full
case, comparisons of the experimental moment of
Inertia of the fluid with some theoretical inertia
solutions for tanks having elliptic and rectangular
sections were made. In the partially full case the
experimentally measured inertias were compared
with the solid inertias for the same tank fullness
with the view of referring the partially full case to
the theoretical solutions for the tank-full condition.
Studies of the effective fluid damping Indicated that
for decaying oscillations the damping is nonlinear
with amplitude. For steady oscillations, the
damping was large as compared with the initial
values for decaying oscillations and tended to in-
:rease with increasing frequency and amplitude.


NACA RM L53E06b

AN EXPERIMENTAL INVESTIGATION OF WHEEL
SPIN-UP DRAG LOADS. Benjamin Milwitzky,
Dean C. Lindquist and Dexter M. Potter. June
1953. 18p. diagrs. (NACA RM L53E06b)


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


NACA RM L53E14a

A FLIGHT INVESTIGATION OF SOME EFFECTS OF
AUTOMATIC CONTROL ON GUST LOADS. Chester
B. Payne. July 1953. 18p. diagrs., tab. (NACA
RM L53E14a)






NACA
RESEARCH ABSTRACTS NO.56 29


The results of a limited flight investigation with a
transport airplane to determine the effects of
automatic control on gust loads indicate that the
loads experienced by the test airplane when
automatically controlled were less than those without
automatic control. The magnitude of the difference
between the loads with and without automatic control
was roughly 7 percent. There was no apparent
change in the effect of the autopilot on gust loads for
a small increase in autopilot sensitivity.


NACA RM L53F12

THE CALCULATED AND EXPERIMENTAL IN-
CREMENTAL LOADS AND MOMENTS PRODUCED
BY SPLIT FLAPS OF VARIOUS SPANS AND SPAN-
WISE LOCATIONS ON A 450 SWEPTBACK WING OF
ASPECT RATIO 8. H. Neale Kelly. September
1953. 35p. diagrs., photo. (NACA RM L53F12)


Experimental data on the incremental lift and pitch-
ing moment produced in the linear angle-of-attack
range by 20-percent-chord split flaps of various
spans and positions on two 450 sweptback wings of
aspect ratio 8.02 have been obtained at a Reynolds
number of 4,000,000 and a Mach number of 0.19 by
pressure-distribution tests in the Langley 19-foot
pressure tunnel. The experimental data have been
compared with calculated increments in span loading
and pitching moment, and the probable causes of the
deviations of the calculated values from the ex-
perimental are discussed.








30


THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM RESTRICTED TO
UNCLASSIFIED, 12. 14/53.




NACA RM A8A21

THE EFFECTIVENESS AT HIGH SPEEDS OF A 10-
PERCENT-CHORD PLAIN TRAILING-EDGE FLAP
ON THE NACA 65-210 AIRFOIL SECTION.
Richard J. Ilk. June 14, 1948. 26p. diagrs., tab.
(NACA RM A8A21)


This report contains the results of a high-speed
wind-tunnel investigation of the effectiveness of a
10-percent-chord plain flap on the NACA 65-210 air-
foil section. The results include an indication of
the lift-producing characteristics and the effective-
ness of the 10-percent-chord flap. From a com-
parison of the characteristics of the 10-percent-
chord flap with those of a 20-percent-chord flap, it
was concluded that, although a reduction in flap-
chord ratio from 0.20 to 0.10 lessens the severity of
the effectiveness loss at supercritical speeds, the
20-percent-chord flap is more effective throughout
the entire range of Mach numbers from 0.3 to 0.875.


NACA RM A8D07

INVESTIGATION OF A THIN WING OF ASPECT
RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND
TUNNEL. I CHARACTERISTICS OF A PLAIN
WING. Ben H. Johnson, Jr. June 2, 1948. 37p.
diagrs., photo. (NACA RM A8D07)


This report presents results of tests of a semispan
model of an unswept wing having an aspect ratio of 4,
a taper ratio of 0.5, and a 4.5-percent-thick diamond
profile. The effects of rounding the profile ridge
were also investigated. The Reynolds number was
varied from 2,000,000 to 10.190,000 at 0.20 Mach
number, and the Mach Number was varied from 0.20
to 0.94 at Reynolds numbers below 3,000,000. No
excessive movement of the aerodynamic center was
observed up to a Mach number of 0.94. The bound-
ary layer separated near the ridge at low Reynolds
numbers, but rounding the ridge diminished this
separation, increased the lift-drag ratio, and de-
creased the minimum drag.


NACA RM A8D20

AERODYNAMIC STUDY OF A WING-FUSELAGE
COMBINATION EMPLOYING A WING SWEPT BACK
630. SUBSONIC MACH AND REYNOLDS NUMBER
EFFECTS ON THE CHARACTERISTICS OF THE
WING AND ON THE EFFECTIVENESS OF AN
ELEVON. Robert M. Reynolds and Donald W.
Smith. October 11, 1948. 56p. diagrs.. photos.,
tab. (NACA RM A8D20)


NACA
RESEARCH ABSTRACTS NO.56

This report presents results of tests of a semispan
model of a wing having its leading edge swept back
630, an aspect ratio of 3.5, a taper ratio of 015,
zero twist, and the NACA 64A006 profile parallellb-.
the airstream. The effectiveness of an elevon as a
longitudinal control was investigated. Aerodynamic
characteristics of the model are presented for a
range of Mach numbers from 0.18 to 0.925 at
Reynolds numbers of 2.35 to 3.55 million and for a
range of Reynolds numbers from 2.35 to 10 million
at 0.18 Mach number.


NACA RM A8E21

TESTS OF A TRIANGULAR WING OF ASPECT
RATIO 2 IN THE AMES 12-FOOT PRESSURE WIND
TUNNEL. III THE EFFECTIVENESS AND HINGE
MOMENTS OF A SKEWED WING-TIP FLAP. Carl
D. Kolbe and Bruce E. Tanting. September 21, 1948.
30p. diagrs., photo. (NACA RM A8E21)


Results of wind-tunnel tests of a semispan model of
a triangular wing of aspect ratio 2 with a skewed
wing-tip flap are presented. Lift, drag, pitching-
moment, and hinge-moment data are included for
subsonic Mach numbers up to 0.95. The flap showed
extremely high hinge moments and low effectiveness
as a longitudinal control. Although less affected by
compressibility, this flap is indicated to be inferior
to a constant-chord flap when applied to this triangu-
lar wing.


NACA RM A8F15

INVESTIGATION OF A THIN WING OF ASPECT
RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND
TUNNEL. II THE EFFECT OF CONSTANT-
CHORD LEADING- AND TRAILING-EDGE FLAPS
ON THE LOW-SPEED CHARACTERISTICS OF THE
WING. Ben H. Johnson, Jr. and Angelo Bandettini.
October 18, 1948. 42p. diagrs., photos., tabs.
(NACA RM A8F15)


Results of tests at low speeds of a thin unswept wing
equipped with a constant-chord leading-edge plain
flap and a constant-chord trailing-edge flap of either
the plain or split type are presented at Reynolds
numbers from 3,000,000 to 10,000,000. The maxi-
mum lift coefficient obtained by an optimum deflec-
tion of these flaps in combination was 1.45 com-
pared to 0.74 for the plain wing with all flaps neutral.



NACA RM A8K19

INVESTIGATION OF A THIN WING OF ASPECT
RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND
TUNNEL. IV THE EFFECT OF A CONSTANT-
CHORD LEADING-EDGE FLAP AT HIGH SUBSONIC
SPEEDS. Ben H. Johnson, Jr. and Verlin D. Reed.
January 20, 1949. 30p. diagrs., photo. (NACA
RM ABK19)







;NACA
RESEARCH ABSTRACTS NO. 56


The report presents results of tests at Mach num-
bers from 0.20 to 0.94 of a thin straight wing of
aspect ratio 4, having a full-span, constant-hord,
leading-edge flap. The wing profile was a modified
diamond having a thickness ratio of 0.042. The in-
vestigation was made at a constant Reynolds number
of 2,000,000. Deflection of the leading-edge flap
produced an increase in mai~mum lift-drag ratio and
an increase in the lift coeificient for maximum lift-
drag ratio for all Mach numbers below 0.94l. At a
Mach number of 0.94, deflection of the flap resulted
.in a decrease of the maximum lift-drag ratio.


NACA RM A8L20

INVESTIGATION OF A THIN STRAIGHT WING OF
ASPECT RATIO 4 BY THE NACA WING-FLOW
METHOD. LIFT AND PITCHING-MOMENT CHAR-
ACTERISTICS OF THE WING ALONE. George A.
Rathert, Jr., Carl M. Hanson and L. Stewart Rolls.
February, 1949. 37lp. diagrs., photos. (NACA
RM A8L20)

This report presents measurements of the hItI and
pitching-moment characteristics of a straight wang of
aspect ratio 4, taper ratio 0.5, having a symmetrical
double-wedge airfoil section with a maxmum thick-
ness of 4.4-percent chord. The tests were conducted
in the Mach number range 0.51 to 1.20, Reynolds
numbers 380,000 to 660,000, by the NACA wing-flow
method. The regents are compared with theory and
with wind-tunnel tests of a similar model. It is
indicated that In the Mach number range 0.82 to 1.00
the model surface, profile, and test Reynolds number
all womad De very Importamr consrlaerarions mn any
attempt to study or predict full-scale characteristics
from small-model tests.


NACA RM ADA24

A PRELIMINARY EXPERIMENTAL INVESTIGATION
OF A SUBMERGED CASCADE INLET. R. Duane
Christiant and Lauros M. Randall. March 25, 149s.
29p. diagrs., photos., 3 tabs. (NACA RM A9A24)



al n ofasberme inlet tnorprtn a cascade
of airlolls for turning and diffusing the entering air.
The geometric variables investigated included ramp

Abld 1ne aoldl d ai an te ,nl andth dute-
center line aft of the cascade of airfoils. The effects
rfthese tarables on the ram-pressure recoveries


NACA RM AOB28

WIND-TUNNEL INVESTIGATION OF TRANSONIC
AILERON FLUTTER. Albert L. Erickson and
Robert L. Mannes. June 1949. 55p. diagrs.,
photos., tab. (NACA RM A9B28)


A new type of aileron flutter is shown to extet in the
transonic speed range. This flutter is characterized
by the fact that only one degree of mechanical free-
dom laI required.


NACA RM A9D08

A FLIGHT INVESTIGATION OF THE EFFECT OF
FLAP DEFLECTION ON HIGH-5PEED
LONGCITUDINAL-CONTROL CHARACTERISTICS.
Maurice D. Whate, Melvin Sadoff, Lawrence A.
Clousing and George E. Cooper. June 30, 1949.
28p. diagrs., photos., tab. (NACA RM A9D08)


Flight tests were conducted on two airplanes to study
the effect of wing flap deflection on longitudinal sta-
bility and control characteristics at supercritical
speeds. Negative flap deflection was beneficial In
one case but not in the other. The relative effects on
longitudinal stability and control of the varianon with
Mach number of angle of attack, airplane tail-off
pitching moment, downwash and dynamic pressure at
the rall, and aileron floating angle are presented.
Attempts to predict the superenticral characteristics
resulting from deflecting the flaps were made, but
the results were inconsastent.


NACA RM A9G20

A PRELIMINARY INVESTIGATION OF THE USE FUL-
NESS OF CAMBER INu OBTAINING FAVORABLE
AIRFOIL-SECTION DRAG CHARACTERISTICS AT
SUPERCRITICAL SPEEDS. Gerald E. Nalzberg,
Stewart M. Crandall and Perry P. Polentz.
October 7,1949. 24p. diagrs. (NACA RM A9G20)


The possibility of employing camber to obtain an air-
foll section which, at moderate hItI coeffacrents and
supercritical speeds, has drag coefficrents signall-
cantly lower than those of the corresponding un-
cambered section as demonstrated. The measured
high-speed sectson characterlstics of an example,
which appeared promising on the baste of calculations
of drag-divergence Mach number, are presented.



NACA RM A9J26

THE EFFECT OF ACCELERATING A HYPOTHETI-
CAL AIRCRAFT THROUGH THE TRANSONIC RANGE
WITIH 4000ROL F .ED. HowarF d(AMMteA J26



Mach n rberhistories of Lh m teon erpr is ed by


cnerls-ue ra n tphd e 8 aslanc cagea Ine Lran-
sonic range obtained by a differential analyzer are
presented. The effect of longstudusal acceleration on
the maximum change in normal acceleration is dis-
cussed. It is shown that the Mach number variation
of the angle of attack for static balance determines to
a great extent the degree of the effect of Longitudinal
acceleration on the normal acceleration. Two
approximate analytical solutions of the longstudlnal
equations of motion are developed and are shown to
compare favorably with the dlfferential-arnalyzer
solutions.


is dis-
cussed. It is shown that the Mach number variation
of the angle of attack for static balance determines to
a great extent the degree of the effect of longitudinal
acceleration on the normal acceleration. Two
approximate analytical solutions of the longitudinal
equations of motion are developed and are shown to
compare favorably with the differential-analyzer
solutions.








NACA
RESEARCH ABSTRACTS NO. 56

NACA RM E8A29

ICE PROTECTION OF TURBOJET ENGINES'BT
INERTIA SEPARATION OF WATER. II ANNetX
SUBMERGED INLETS. Uwrevon Glahn. June 6,
1948. 21p. dialgre., photos., tab. (NACA
RM E8A29)

Aerodynamic and icing studies were conducted on a
one-half-acale model of an annular submerged inlet
for use with axial-flow turbojet engines. Pressure
recoveries, screen radial-velocity profiles, circum-
ferential mass-flow variations, and icing character-
istics were determined at the compressor Inlet. In
order to be effective in maintaining water-free indue
tion air, the inlet gap must be extremely small and
ram-pressure recoveries consequently are low, the
highest achieved being 65 percent at inlet-velocity
ratio of 0.86. All inlets exhibited considerable
screen icing. Severe mass-fLowr shiftsl occurred at
angles of attack.

NACA RM E8B27

ANALYSIS OF PARAMlETERS FOR THRIWT CON-
TROL OF A TURBOJET ENGINE EQUIPPED WITH
AIR-INLET THROTTLE AND VARIABLE-AREA
EXIIAWT NOZZLE. Aaron 8. Bokeenbom and
Melvin 8. Feder. August 10, 148. 62p. diagre.
(NACA RM E8B27)


By means of an anlyIsIII developed in this report,
characteristics of a typical turbojet engine equipped
with an air-inlet throttle, a variable-area exhaual
nozzrle, or with a combination of the two were pre-
dicted and control-parameter curves were chosen
from which such control systems can be designed.
The thrust control characteristics of a turbojet en-
gine could be substantially improved by the addition
to the conventional control system of either an air-
Inlet throtle, a variable-area exhanust mosale, or a
combination of the two.

NACA RMd EgCOg

EFFECT OF INLET-AIR PARAMETERS8 ON COMl-
BWSTION LIMIT AND FLAME LENGTH IN 8-plCH
DIAMETER RAM-JET COMBWBTION CHAMBER.
A. J. Cervenka and R. C. Miller. July 2, 1948.
24p. diagra. (NACA RM E8COp)


The effect ofilet-r pan mterra on combustion
diameter ram-jet comtmation chamber. For a given
combustion-chamber length, combustion efficiency
increased with increasing inlet-air temperature or
dceeng ine-air velocity and prasue Fae

inlet-air pressure, inlet-air velocity, and decreasing
inlet-air temperature.

NACA RMa E8018

PRELMNARY RESULTS OF NATURAL ICING OF
AN AXIAL-FLOW TURBOJET ENGINE. Loren W.
Acker. August 8, 1948. 16p. diagrs., phatog.
(NACA RM E8C18)


A fligh investigation in natural kcing conitions was
conducted to determine effect of ice formations on
performance of an artal-flow turbojet eaglue. Re-


32

NACA RM A9K01

AN ANALYSIS OF THE FORCES AND PRESSURE
DISTRIBUTION ON A WING WITH THE LEADING
EDGE SWEPT BACK 37.25o. George G. Edwards
and Frederick W.Boltz. March30, 1950. 102p.
diargrs., photos., 2 tabs. (NACA RM A9K01)


Results of wind-tunnel tests are presented for a semi-
span model of a wing having the leading edge swept
back 37.250, aspect ratio 6, taper ratio 0.5, zero
twist, and NACA 641-212 sections normal to the
quarter-chord lane of these sections. Lift, dra8.
pitching moment, and pressure-dletrlbution data
were obtained for a range of Mach numbers from 0.18
to 0.94, at a Reynolds number of 2,000,000. Similar
data were obtained for Reynolds numbers of
1,100,000, 3,000,000, and with roughness applied to
the wing (at a Reynolds number of 2,000,000). An
analysis is made to correlate the changes in pressure
distribution over the wing with the changes in the
total forces.




NACA RMd E8A27

ICE PROTECTION OF TURBOJET ENGINES BY
INERTIA SEPARATION OF WATER. I -
ALTERNATE-DUCT SYSTEM. Uwre von G~lan.
May 28, 1948. 77p. dliagr., photos., 3 tabs. (NACA
RM E8A27)

Aerodynamic and icing investigations of internal
water-inertia separation inlets designed to prevent
automatically entrance of large quantities of water
into a turbojet engine in scing conditions was con-
ducted on a one-half scale model. A simplified
analytical approach to the deengn of internal wataer-
Inertia separation inlet is included. Results show
that in order to be effective In preventing screen and
guide-vane icing for an inlet of this type, a ram-
pressure recovery of 75 percent was attained atl de-
eign inlet-velocity ratio in an Icing condition. For
nonicing operation, ram-preseure recovery is
comparable to direct-ram inlet.


NACA RM E8A28

ICE PROTECTION OF TURBOJET ENGINES BY
INRIA BE ART ONMOF WATER. G SING E

1948. Ulp. diagre., photos (NACA RM E8A28)


Investigation of a single-offset-duct system designed
to prevent entrance of water anto a turbojet engine
was conducted on a half -scale naelle model. An in-
vestzgation was made to determine ram-pressure
recovery and radial velocity profiles at the com-
pressor section and scing characteristics of such a
duct system. At a design inlet velocity of 0.77, the
maximum ram-pressure recovery attained with
effective water-separatmng Inlet was 77 percent,
which as considerably less than attainable with a
direct-ram Inlet. Continuous heating of the
accessory-housing surface would be required for
Inlets that have a small ice storage space.







NACA
RESEARCH ABSTRACTS NO. 56

aults are presented for a fligh in which the icing
rate varied from 5.1 to 3.1 inches per hour. During
45 minutes in icing, tail-pipe temperature increased
from 761o to 10508 F and jet thrust decreased from
1234 to 910 pounds. cet penetrated to the seceond-
stage stator blades.
NACA RM EBDIS

EXPERIMENTAL INVESTIGATION OF HOT-GAS
BLEEDBACK FOR ICE PROTECTION OF TURBO-
JET ENGINES. I NACELLE WITH OFFSET AIR
INLET. Edmund E. Callaghan, Robert S. Rugged
and Richard P. Krebs. July 8. 1948. 22p. diagrs.
(NACA RMl E8D13)


Investigations were conducted on a two-thirds-scale
model of a turbojet-engine nacelle to provide icing
protection for the engine, protective screen, and
accessory housing by the introduction of hot gasee
into the inlet-air stream. The hot gas was injected
into the intet-air stream from a series of orifice
holes around the Inlet. Results indicate that If at
minimum kinetic temperature of 32o F to maintained
througout the screen, adequate Ice protection will
be obtained for the entire Installation.

NACA RMd E87018

NATURAL ICING OF AN AXIAL-FLOW TURBOJET
ENGINE IN FLGHT FOR A SINGLE ICING CONDI-
TION. Isren W. Acer. August12, 1948. 16p.
diagra., photos. (NACA RM E8F01a)


A flight investigation In natural icing conditions was
conducted to determine the effect of ice formations
on performance of axial-flow turbojet engines. Re-
salte are presented for a f~lgt in which liquid-water
content varied from 0.077 to 0.400 gram per cubic
meter. During 60 minutes in icing, tail-pipe temper
sture Increased from 86150 to 965o F and the jet
thrust decreased Imom 1950 to 1700 pounds. The en-
gIne was satisfactortly accelerated to take-off power
near the end of the icing period.


NACA RM E8J25b

ANALYSIS OF EFFECTS OF INLET PRESSURE
LOBBES ON PERFORMANCE OF AXIAL-FLOW
TYPE TURBOJEGT ENGINE. NewelL D. Sanders and
John Palasics. November 4, 1948. 39p. diagrs.
(NACA RM E8J25b)


The effects of Inlet pressure losses on the net thrust
and specific fuel consumption for an axial-flow type
turbojet engine have been determined. At Low alti-
tudes and flight Masch numbers, a 10-percent lose in
Inlet total pressure will result in a 14-percent lose
in net thrust and a 15-percent increase in specific
fuel consumption, and at high altitudes and f~ligt
Mach numbers a 10-percent Inlet total-pressure lose
produces a 22-percent lose in net thrust and an In.
crease of 16 percent in specific fuel consumption.


33

NACA RM EBJ25c

INLET ICING AND EFFECTIVENESS OF HOT-GAS
BLEEDBACK FOR ICE PROTECTION OF TURBO-
JET ENGINE. William A. Fleming and Martin J.
Saart. November, 1498. 37p. dligre., photoo.
(NACA RM E8J25c)


A method of protecting the inlet of a turbojet engine
from Ice was studied in the NACA Cleveland altitude
wind tunnel whereby hot gas was bled from the tur-
blne Inlet and Injected into the induction system
ahead of the compreesor. At ambient-air tempera-
tures of 250 to 300 F, the hot-gasl bleedback system
removed Ice from the compressor inlet in 3 to 5
minutes; however, at ambient temperatures of about
00 F, Ace formed on the compressor inlet very
slowly below 11,000 rpm with the hot-gas bleedback
system in operation.

NACA RM EOB23

EFFECT OF H(Fr-GAS BLEEDBACK ICE PREVEN-
TION ON PERFORMANCE OF A TURBOJET EN-
GINE WITH FlKED-AREA TAIL-PIPE NOZZLE.
Robert 0. Diets, Jr. and Richard P. Krebe.
May 16,194.9. 2p. diagra. (NACA RM EOB23)


An analytical Investigation was conducted to show
that the Inlet of a turbojet engine can be protected
from ice accretions by bleeding hot gases from other
locations within the engine to the Inlet without undue
lose in thrust. Bleedback from the combustion cham-
ber was preferable to tail-pipe bleedback because the
pressure was greater, less bleedbackr was required,
and smaller thrust losses resulted. The thrust avail-
able at take-off from an engine protected against
icing conditions in temperatures as Low as -40o F
with liquid-water contents as high as 2.5 grama per
cubic meter exceeded the thrust available from the
same engine in an amblent-air temperature of 1000 F.

NACA RM ESC16.

EXPERIMENTAL INVESTIGATION OF HCYT-GAS
BLEEDBACK FOR ICE PROTECTION OF TURBO-
JET ENGINES. II NACELLE WITH LONG
STRAIGHT AIR INLET. Edmund E. Callaghan and
Robert 8. Ruggedl. May 26, 1949. 34p. diagre.,
photo. (NACA RM E9C16)


Investigations were conducted on a two-thirde-scale
model of a turbojet-engine nacelle to provide basic
design criterions for hot-gas bleedback eyetems.
The hot gas was infected into the inlet-air stream
from a sortee of orifices around the Inlet. General
rules for obtaming a satisfactory orifice configurar-
tion are presented. Satisfactory agreement between
calculated and measured heat requirements was ob-
tained.

NACA RM E9E06

ANALYSIS AND PRELIMINARY INVESTIGATION OF
EDDY-CURRENT HEATING FOR ICING PROTEC-
TION OF AXIAL-FLOW-COMPRESSOR BLADES.
Thomas Dallas and C. E~laman. August 8, 1949.
66p. diagrs., photos., tab. (NACA RM E9E0)







NACA
RESEARCH ABSTRACTS NO. 56

blne blade temperature during periods of accelera*
tion and deceleration and that the time required for
rotor speed to reach equilibrium was appgm~a~tely
the same for accelerations and dcesle~orationslq
same magnitude for the same range of speeds.

NACA RM LBASOa

A COMPILATION OF THE PRESSURES MBABSURED
ON A WING AND AILBRON WITH VARWIWt
AMOUNTS OF SJWBBP IN THIE LANGLEY8-P00@
HIGH-8PEED TUNNEL. Richard T. Whitcomb.
April IS, 1948. 84p. diagra., 18 tabs. (NACA
RM L8A30a)


A completion La made in tabular form al all the
pr au mw 3re o t b ellb l
sweepback and sweeplorward at high subsoni Mach
numbers in the anogley 8-root ghl-~peed tunnel.

NACA RM L8CHl

EdDNGITUDINAL STABILITY ANDm CONTRDL CRAR-
ACTEIGSTICS AT TRANBONIC BPEEDS OF A
auMISPA AIRPLANE MODEL HAVRIS A IbO
IWILDHACEr WRIN AND TAll. AY OWalgADlRbT
TH TSAMONIC-BtlMPP MBTRID. M. Lqe
sp ~"a June 4 seas as.dhp, e
tab. (NACA RM LACll)


Tests were made usin the transonic-bump method to
determine the stability and control charaectrlatica in
the transonic range of a semispan airplane model
having a go0 aweptback wing and tail. The Ilt.t drag,
and pitching-moment coaeffcints for two sabiliser
settings were obtained at various angles of attack
through a Mach number range from 0.50 to 1.23.

NACA RMY LAC12

AN INVESTIGATION OF THE DOWNWASHR BEHIND
A HIGH-ASPECT-RATIO WING WITH VARIOUS
AMOUN'Il OF SWEEP IN THE LANGLEY 8400T
HIGH-SPEED TUNNEL. Richard T. Whitcomb.
May 11, 1948. 25p. diagra., 2 tabs. (NACA
RM L8C12)

Downwash angles have been measured at verrtical
poaltions at the probable tail location behind a high-
aspect-ratio wring with an NACA 65-210 section with
no sweep and 300 and 45o of sweepback and sweeplor-
warrd in conjunction with a fuselage at Mach numbers
up to 0.96. Presented are the variations of down-
wash angles wvith normal-force coeficient and Mach
number.
NACA RM L8C15

LANGLEY FULL-SCALE-TUNNEL INVESTIGATION
OF THE CHARACTERISTICS IN YAW OF A
TRAPEZOIDAL WING OF ASPECT RATIO 4 WITH
CIRCULAR-ARC AIRFOIL SECTIONS. Ralph W.
May, Jr. and George l. Stevens. August 30, 1948B.
19p. diagrs., photos. (NACA RM L8Cl5)


Results are given of an investigation at an approdL-
mate Reynolds number of 4,.100,000 and March number


An analysts and a preliminary experimental investi-
gation of the use of eddy currents for heating the
blades of an azlal-flow compressor of a turbojet en-
gine as a means of ice protection is presented*
Steinmets's analyale of eddy currents in flat plates is
extended and formulas are derived that permit calmu-
lations of currents, voltages, flux distribution, and
pDWer When Operation is not limited to conditions of
constant permeability. The application of eddy-
current heating appears to be feasible for the genlera-
ELon of power densities required for Icing protection
of adal-flow-compressor blades.

NACA RM EgBE12

EXPERIMENTAL INVESTIGATION OF HOT-GAS
BEDAC O ICEA POECTO OFOT B-

RTRA~t0HTAIR INLET. Robert 8. Roggert and
Edmound EACA~~bn 981 st 4, 19. 4p. diagram ,


An aerodynamic and Scing investigation was conducted
on a two-thirds-scale model of a turbojet-engine
narcelle utilizing a hot-gas bleedback system for Sce
prevention. The optimum temperature distribution
aspMe the model was obtataed as a bleedback of 4.0
Pagage fas a gas temperature el SA0ll P and was n-
depndwlof teamet velocity. lMEs psetection fo te M
aceseEM P house could not as M~L auinssaesas
lcess than 6.0 percent for a gas temperature of
1000P F. Sattefactory agreement between measured
and calculated heat requirements for Icing conditions
was obtained.

NACA RM EGI27

ALTIUDE-CHAMBER PERFORMANCE OF
BRITISH ROLLS-ROYCE NENE II ENGINE. II -
18.41-INCHDIAMETER JET NOZZLE. J. C.
Armatrong, H. D. Wilated and K. R. Vincent.
October 26, 1949. 55p. diagrs., 2 tabs. (NACA
RM E9I27)

An altitude-chamber investigation of British Rolla-
Royce Nene II turbojet engine was conducted over
range of altitudes from sea level to 60,000 feet and
ram-pressure ration from 1.00 to 3.50, using 18.41-
inch-diameter jet nozzle. The 18.41-inch-diamseter
jet nozzle showed a small decrease in net-thrust
specific fuel consumption from the standard 18.75-
inch-diameter jet nozzle at cruise conditions.

NACA RM E9K25a

TRANSIENT OPERATING CHlARACTERISTICS OF A
TURBOJET ENGINE WHEN SUBJECTED TO STEP
CHANGES IN FUEL FIDW. Arthur H. Bell and J.
Elmo Farmer. February20, 1950. 45p. diagra.,
photos.. 2 tabs. (NACA RM EglK25a)


An investigation was conducted on a typleal centrifu-
ga compressor type turbojet engine to determine
operating characteristics during changes in engine
speed resulting from approximate step changes In
fuel flowr. The data show the variation with time of
engine speed, thrust, fuel flow, and critical gas and
material temperatures as caused by sudden changes
in fuel fowr. Results show that the gas temperature
at the tail-cone exit indicated the approximate tur-






NACA
RESEARCH ASSTRACTS NO. 56

of 0.07 to determine the characteristics in yaw of a
trape~oidal wing of aspect ratio 4 with circular-arc
airfoil sections and 0.30-chord rear and drooped-
nose flaps. Measurements of the lift, the lateral
force, the rolling moment, and the ynawing moment of
five configurations were made through at range of
angle of attack at several yaw angles ranging from
approximately -60 to 180'

NACA RM LAD06

EFFECT OF HIGH-LIFT DEVICES ON THE LONGI-
TUADINAL AND LATERAL CHARACTERISTICS OF A
45o SIWEPTBACK WING WITH SYMMETRICAL
CIRCULAR-ARC SECTIONS. Eugene R. Ouryanskry
and Stanley Lipeon. October 1, 1948. 45p. diagra.,
photo. (NACA RM L8D06)


An investigation was made in the Langley full-scale
tunnel at a Rey~noldsl number of 4.5 x 10" of the longi-
tudinal characteristics of several leading-edge and
tralling-edge flap configurations and the lateral char-
acteristics of one flapped configuration of a 45o
sweptback wing having symmetrical circular-are
sections, aspect ratio 3.5, and taper ratio 0.5. In-
veslltigton included tests of chordwise fences with
and without a rounded leading-edge modification in-
stalled on the outer semispan of the wing.

NACA RM L8DlI4

PRELIMINARY CORRELATION OF THE EFFECT OF
COM6PREgSIBILITYY ON THE6 VACATION OF THE
SECTION AERODYNAMIC CENTER AT SUB-
CRITICAL SPEED. Edward C. Polhamus.
November 3, 194. Tp. diagr., tab. (NACA
RM L8D14)

A correlation of available two-dimensional airfoil
data to determine the effects of compressibility on
the location of the section aerodynamic center at low
111 coefficients has been made. The results indicate
that large forward or rearward movements of the
aerodynapmic center with Mach number are possible.
It appears that tlhickess ratio is an important
parameter controlling the variation with Mach num-
ber.
NACA RM L8D30

EFFECT OF LEADING-EDGE HIGH-LIFT DEVICES
AND SPLIT FLAPS ON THE MAXIMUM(-LIFT AND
LATERAL CHARACTERISTICS OF A RECTANGU.
LAR WING OF ASPECT RATIO 3.4 WITH
CIRCULAR-ARC AIRFOIL SECTIONS AT REYNOLDS
NUMBERS FROM 2.9 x 106 TO 8.4 x 106. Roy H.
Lange and Ralph W. May, Jr. November 10, 1948.
70p. diagre., photos. (NACA RM L8D30)


Results are presented of an investigation at high
Reynolds numbers and Low Mach numbers in the
Langley full-scale tunnel to determine the effects of
leading-edge high-lift devices and split flape on the
maximum-lift and lateral characteristics of a rectan-
gular wing of aspect ratio 3.4 with circular-arc air-
foil sections. Measurements were made of the static..
aerodynamic characteristics in pitch and in yaw of the
Ibasic wing and of the wing with several leading-edge


35

high-lift devices and 0.20-hord split flape deflected
alone and in combination with one another. Scale
effects were Investigated at Raynolds numbers rang-
Ing from 2.9 x 106 to 6.4 x 106,

NACA RM L8Fl6

LOW-SPEED PRESSURE DISTRIBUTIONS OVER THE
DROOPED-NOSE FLAP OF A 42o SWEPTBACK
WING WITH CIRCULAR-ARC AIRFOIL SECTIONS
AT A REYNOLDS NUMBER OF 5.3 x 10 Stanley H.
Spooner and Robert L. Woods. September 23, 1948.
28p. diagrs., tab. (NACA RM L8F16)


Pressure distributions over the drooped-nose flap of
a 42o sweptback wing having circular-arc airfoil
sections are presented for several angles of attack.
The drooped-nose-flap normal-force and hinge-
moment coeffaclents are determined. The effects of
nose-flap deflection and span and of half-span
trailing-edge splat flaps are shown. The Reynolds
number and Mach number of the investigation were
5.3 x 106 and 0.11, respectively.

NACA RM L8F17

INVESTIGATION OF FLOW CONDITIONS AND THE
NATURE OF THE WALL-CONSTRICTION EFFECT
NEAR AND AT CHOKING BY MEANS OF THE
HYDRAULIC ANALOGY. Clarence W. Matthews and
Ray H. Wriht. September 1, 19.8. 6p. diagre.
(NACA RMI LBFIT)


Field surveyed are presented of subcritical, super-
critical, and choking flow about several airfoils at
zero lift in a water channel. A qualitative study is
made of the effects of the channel walls on super-
critical flow about the airfoils, of the development of
the choking condition from sup~ercritical flow, and of
the effects of airfoil characteristics on the choking
stream Mach number. The experimental wall-
interference results are compared with results ob-
tained by theoretical calculations.
NACA RM L8F29

AN INVESTIGATION AT LOW SPEED OF A 51.30
SWEPTBACKE SEMIB(PAN WING WITH A RAKED TIP
AND W~ITH 161.7-PERCENT-CHORD AILERONS
HAVING THREE SPANS AND THREE TRAILING-
EDGE ANOLES. Jack Fischel and Leslie E.
Schnetter. July 21, 1948. 52p. diagrer. photos., 3
tabs. (NACA RM L8F09)


Contained results and discusalon of aerodynamic and
lateral-control Inveesltigton at low speed of a 51.30
aweptback semispan wing with a raked tLp and with
16.71-percent-chord sealed plain allmern having
spans of 34, 66, and 9S percent of the apan of a full-
span aileron, each with tralling-edge angles of So,
1o, and 25o. Lift and lateral-control data were ob-
tained throughout a large angle-of-attack range.
Spanwisle variations of rolling-moment and hinge-
moment parameters, as well as the effects of the
aileron trailing-edge angles, were determined from
the arlleron configurations tested.







NACA
RESEARCH ABSTRACTS NO. 56

NACA RM L8IO3

THE EFFECTS OF HIGH-LIFT DEVICES OY THE
LOW-SPEED STABILITY CHARACTERISTICS OB~ s
TAPERED 37.50 SWEPTBACK WING OF ASPECT
RATIO 3 IN STRAIGHT AND ROLLING FLOW.
M. J. Queijo and Jacob H. Lichtenstein.
November 9, 1948. 27D. diagrs., photo., tab.
(NACA RM L8103)


Contains results of tunnel tests to determine effects
of various combination of sphlt flaps, state, and
nose alats on the stalbilrty enaracteristice of a
tapered 37.50 sweptback wing of aspect ratio 3 to
straight and rolling flow.

NACA RM L610)

TRANSONIC-FLUTTER INVESTIGATION OF WING
ATTACRIED 'IC TWO IA)W-ACCELERATION
ROCKET-PROPELLED VEHICLES. Reginald R.
Lundstrom, William T. Lauten, Jr. and Ellwryn E.
Angle. November, 198. 24p. diagrs., photosl.
3 tabs. (NACA RM L810)


Two low-acceleration transonic-flutter vehicles were
launched and flown. The first carried two test wings,
one of which flutter ad at M = 0.92 at a frequeny of
61.4l cycles per second. The reference Rutter speed
determined from two-dimensional theory for an un-
swept wing in incompressible flow Is conservative
when compared to the experimental flutter speed. The
second vehicle caused two test wings, one of which
failed at M = 0.71 becuase of a low-frequency diver-
gent oscillation. Since this failure was not caused by
conventional flexure-tors on flutter, no comparison
with a reference flutter speed can be made.

NACA RM L8J11

INITIAL EXPERIMENTS ON FLUTTER OF UN-
SWEPT CANTILEVER WINGS AT MACH NUMBER
1.3. W. J. Tuovila, John E. Bakerand Arthur A
Regler. Janualry 6, 194. 23p. diagrs.,photos., 2
labs. (NACA RM L8J11)


This paper contains experimental data for the bend-
ing toralon flutter of unswept cantilever wings at
Mach number 1.3. Comparison with the theory of
flutter at sulpersonic speeds indicates that the theory
la applicable in predicting cantilever-wilng flutter.
These preliminary experiments consider also various
overall effects of center-of-gravity location, elastic-
axis location, mass-density parameter, section
shape, and some wing-tip moments of Inertia.
NACA RM L8Kl2

LATERAL-CONTROL INVESTIGATION ON A 3'?'
SWEPTBACK WING OF ASPECT RATID 6I AT A
REYNOLDS NUMBER OF 6,800000. Robert R.
Graham and W~lmm Koven. January 9T, 1949.
58p. diagra., photo. (NACA RM1 LBKl2)


Lateral-control devices including an alleron, a plain
spoiler at two chordwlae Locations, and a step


36

NACA RM L8G05

FULL-SCALE INVESTIGATION OF AN EQUI-
LATERAL TRIANGULAR WING HAVING 10-
PERCENT-THICK BICONVER AIRFOIL SECTIONS.
Edward F. Whittle, Jr. and J. Calvin Lovell.
September30, 1948B. 32p. diagre.,photos. (NACA
RM L8005)


Contains the results of an investigation made at a
Reynolds number of 6.0 x 106 to determine the low-
speed characteristics of a wing having triangular
plan form, 600 of sweepback at the leading edge, and
10-percent-thick biconvex airfoil sections. The In-
vestigation consisted of the determination of the ef-
fects on the stalhng and static longitudinal charac-
terlsuecs of semispan and full-span leading-edge and
trailtng-edge flaps and the effects on the static
lateral stability characteristics of a vertical fan.

NACA RMd L8G22

DOWNWASR, SIDEWASH, AND WAKE SURVEYS BE-
HIND A 42o SWEPT CK WING AT A REYNOLDS
NUMdBER OF 6.8 x 10" WITH AND WITHOUT A
SIIMULATED GROUND. G. Chester Furlong and
Thomas V. Ballech. December 3, 1948. 77p.
diagrs., photos., tab. (NACA RMl L8G22)


An investigation, with and without a simulated
ground, has been conducted to provide flow inclinr-
thon and wake data behind a 42o sweptback wing.
Tests were made for two model configuration;
namely, the plain wing and the wing with Inboard
trailing-edge split flape and outboard leading-edge
flaps deflected. Contour charts of downwash, side-
wash, and dynamic -pressure ratto at two longitudinal
stations behind the wing (2.0 and 2.8 mean aero-
dynamic chords) are presented. Integrations have
been made to obtain variations of average downwash
and dynamic-preeaure ratio with angle of attack.r)
The possibility of extending the Ultng-line method '
used for calculating the downwash behind unswept
wings to the case of a sweptback wing has been
briefly investigated.

NACA RM L8H20

AN INVESTIGATION AT LOW SPEED OF A 51.3o
SWEPTBACK SEMISPAN WING EQUIPPED WITH
16.71-PERCENT-CHORD PLAIN FLAPS AND
AILERONS HAVING VARIOUS SPANS AND THREE
TRAILING-EDGE ANGLES. Jack Fischel and Leslie
E. Schneiter. November 12, 1948. 81p. diagrs.,
photo., 2 tabs. (NACA RM L8H20)


Contains results and discussion of an aerodynamme
inveetagation at low speed of a 51.30 sweptback semi-
epan wing equipped with 16.7-percent-chord plain
flaps and ailerons having various spans, spanwlae
locations, and trailing-edge angles. Variabion of the
lift and lateral-control data with flap or alleron span
and spanwrise location were determined through a
large angle-of-attack range. The effects of aileron
tralling-edge angle and the aerodynamic characterle-
ticse ofa speller- iledFon configuration on the wing







NACA
RESEARCH ABSTRACTS NO. 56

spoiler were investigated. The effects of stall con-
trol and hgh-lift devices including a leading-edge
flap, slat, and drooped leading edge and split and
double slotted trailing-edge flape on the lateral con-
trol characteristics were also investigated. Force
and moment data are presented for all the devices
and in addition hinge-moment and balance compart-
ment pressure coefficients are presented for the
alleron.

NACA RMd L8L01

ANALYTICAL CONSIDERATIONS REGARDING A
CONTROL-SURFACE BALANCE ARRANGEMENT
FOR SUBSONIC AND SUPERSONIC FLIGHT.
Thomas A. Toll andGlenn R. Adair. January 10,
1949. 20p. diagra. (NACA RM L8L01)


Presents results of a simplinied analyrsis to indicate
the possibility of utilizing the principle of the sealed
internal balance to reduce control-surface hinge
moments at both subsonic and supersonic apeeds.
Certain designl considerations and possible
advantages of such a balance are discussed.

NACA RM L8L29

LOW-SPEED STATIC-STABILITY AND ROLLING
CHARACTERISTICS OF LOW-ASPECT-RATIO
WING OF TRIANGULAR AND MODIFIED TRIANGIN
LAR PLAN FORMS. Byron M. Jaquet and Jack D.
Brewer. March 29, 1949.. 44p. diagrs., photos.,
tab. (NACA RM L8L29)


Presents results of an investigation conducted in the
Langley stability tunnel to determine effects of
changes in profile and aspect ratio on the low-apeed
static-stability and rolling characteristics of trian-
gular wings. The investigation was extended to
determine the effects of adding fans to the upper sur-
face and of cutting portions from the tips of a trian-
gular wing to form low-aspect-ratio tapered wings.
Results are compared with available swept-wrng
theory and triangular-wang theory.

NACA RM L9B5a a

COMPARISON OF SEMISPAN DATA OBTAINED IN
THE LANGLEY TWO-DIMENSIONAL LOW-
TURBULENCE PRESSURE TUNNEL AND FULL-
SPAN DATA OBTAINED IN THE LANGLEY 19-
I*DOT PRESSURE TUNNEL FOR A WING WITH 4100
SWEEPBACK OF THE 0.27-CHORD LINE. Jonee
F. Cahll. April 22,1949. 33p. diagre., photo.
(NACA RM L9B25a)


Results obtained in this investigation show that data
obtained from semispan-waing tests in the Langley
two-dimensional low-turbulence pressure tunnel may
be expected to be in good agreement with data ob-
tained from full-span tests of the wing alone except
for unusually sensitive configurations where the lift
distribution is such that small disturbances may
cause a radical change in the Location of the original
stall or In the manner In which the stall progresses.


37
NACA RM LAB26b

CONTINUATION OF WING FLUTTER INVESTIGA-
TION IN THE TRAle0)NIC RANGE AND PRESENT*
TIONI OF A LIMITED SUMMARY OF FLUTTER
DATA. William T. Lauten, Jr. nd J. G. Barmby.
April 21, 1949. 21p. diagr., photos., 2 tabs.
(NACA RM: L9B25b)


Results of four unawept flutter airfolsa attached to
two freely falling bodies are reported. One airfoll
fluttered at a transonic Mach number of I.17.
Fltter frequency and phasingnwere re ordd A
ssersonic flutter data is Included and a critical
lutter region is defined.

NACA RM L9E18

EFFECTS OF VARIOUS OUTBOARD AND CENTRAL
FINS ON LOW-8PEED STATIC-STABILITY AND
ROLLING CHARACTERISTICS OF A TIRIANGULAR-
WING MODEL. Byron M. Jaquet and Jack D.
Brewer. June 29,1949. 60p. dlagra., photoo.,2
tabs. (NACA RM L9E18)



RilTy tus e eto temn elet oe f va ious utboa d
and central-fin arrangements on low-speed statrc-
stability and rolling characternetwos of a triangular-
wing model are presented. One of the purposes was
the determination of the fan configuration to provide
good directional stabulaty throughout the Inft-
coefficrent range.

NACA RM L9F24

EXPERIMENTAL INVESTIGATION OF THE EF-
FECTS OF SWEEPBACK ON THE FLUTTER OF A
UNIFORM CANTILEVER WING WITH A VARIABLE
LOCATED CONCENTRATED MASS. Herbert C.
Nelson and John E. Tomassoni. August 31, 1949.
33p. diagrs., tab. (NACA RM L9F24)


An experimental investigation of the effect of swreep-
back on the flutter characteristics of an untapered
cantilever wing carrying a concentrated wreight is re-
ported. The Investigation covered sweepback angles
of Oo, 450, and 600. The weight used in the Investi-
gation was approximately 14l percent heavier than the
wing. The data presented may be used in conjunction
with analytzeal methods of predicting flutter to indi-
cate the validity of the method used.

NACA RM 09817

UDW-SPEED STATIC IDNGITUDINAL STABILITY
CHARACTERISTICS OF A CANARD MODEL HAVING
A 600 TRIANGULAR WING AND HORIZONTAL TAIL.
William R. Bates. November 9, 1949. 1p. diagre.,
tab. (NACA RM L9HI7)


Contains results of force tested and flow surveys made
in the Langley free-flight tunnel to determine the
stability and control characteristics of a canard
model having a 600 triangular plan-form wing and
horizontal tail. Tests were made with the horlEOntal
tail used as a fixed-nose elevator or floating freely
at various tab deflections.







NACA
RESEARCH ABSTRACTS NO. 56

The results of an investltigaio of the pressure dis-
tirintlan over the full-spn droop-noe flap of a wingr
with the leading edge swept back 47.50 and having
symmetricleP circulr-are aidioll sections are'pg~-
sented in thle pper. The configurations tested ias
eluded the basic configuration, the full-upan droop-
nose flap deflected 100, 200, 300, and 40D, the eami-
span plain flpp deflected 400, and both flaps deflected
400. The investigation was conducted in the LasngleF
full-scale tunel at a Raynolds number of 4.3 a 106
and a Mach number of 0.01.




PRELIMINARY EX~PERIMENTBL ON PDHCES AND
MIOMENTS OF AN OSCILLATING WING AT HI~Gl-
SUBSONIC SPEEDS. 8. A. Clevelnson nd B.
Widmayer, Jr. February 20, 1950. 8p. diagr.,
tab. (NACA RULM LK2


Experimental results of Hifts and moments about the
Quarter chord of an oscIllatng wing at high-subsoni
Mach number are presented. A comparison of the
experimental manitude of the lft rector with the
theory asI given by Dietze showed good Pagremeat.
Comparisons of the moments and the quadratue cnm
ponent of IR lftwh theory indicated that some reflee-
ments in the testing technique are necessary for the
emprimental determination of these untitlees la te


NACA~ RMLOLOP

THE EFFECT OF AILERON SPAN AND SPANWISE
IA)CATION ON THE LOW-SPEED LATERAL CON-
TROL CHARACTERISTICS OF AN UNT~APERED
WING OF ASPECT RATIO 2.00 AND 450 SIWEEP-
BACK. Rodger L. Naeaeth and William Ml. O'Hare.
February 10,1950. 20p. diagra. (NACA
RM L9LOOa)


P esnt rselt of lor-pe a MM d-unne latd

0 5S-chord flap-typ alleor on the latteral control

asp ct ratto 2 009 4n50 sae cbel The tests weEW



MACA RM L9Ll3

ANALYSIS OF V-g DATA OBTAINED FROM~d
SEVERAL NAVAL AIRPLANES. James O. Thornton,
July, 1950. 28p. diagrs., 4 tabs. (NACA
RM 09L13)


V-g records obtained with the F8F-1, BB2C-5, and
F6F-5 airplanes from 1945 to 199 are analysed
statistically to give the frequency of maldmum lands
and airspeeds made in bombing, gunnery, dive pull-
outs, and simulated attacks. Compoolte V-g dia-
grame are shown for the FIU-4 and TBM-3 air-
planes.


38

NACA RM LDHIG

LATfERAL-CONJTROL INVESTIGATION AT A
REYNOLDS NUMBER OF 5,300,000 OF A WING OF
ASPECT RATIO 5.8 SlWEPTFORUWARD 32O AT THIE
LEADING EDGE. Robert R. Graham. February 7,
1950. 4p. diagre.,photo. (NACA RMIL9HIS)
Lateral-control devices, including an alleron and
plain and step spoilers, were investigaed. The ef-
fects of an inboard leading-edge flap and a trailing-
edge double-alotted flap on the lateral-control char-
acteristics were also investigated. Force and
moment data are presented for all the devices and,
in addition, hinge-moment, normal-force, and
balance -compartment -pressure coefficients are pre-
sented for the aileron.

NACA RM L9J07

LOW-SPEED PRESSURE-DISTRIBUTION AND FLOW
INVESTIGATION FOR A LARGE PITCH AND YAW
RANGE OF THREE LOW-ASPECT-RATIO POINTED
WINGS HAVING LEADING EDGE SWEPT BACK 600
AND BICONVEX SECTIONS. Ralph W. May, Jr. and
John G. Hawes. November18, 1949s. 109p. diagre.,
photos., tab. (NACA RM L9J7)


Low-speed pressure distributions and flow character-
Istics through a yaw range of 00 to 350 and an angle-
of-attack range through the stall are presented for
three small-scale Lowr-aspect ratio pointed wings
having 10-percent-thick biconvex sections, 600 swept-
back Leading edge, and 00, 300, and -300 trailing-
edge sweep. Also given are section Ilft coefficients
and centers of pressure at zero yaw, and wing span-
wise load distributsons, force, and moment coeffi-
clents throughout the yaw range. A correlation is
established between the pressure distributions and
the vortex flow characteristics.


NACA RM LS.I2a


""FCRS EON RLO NREIRAITN NO FTHE FLUTTER
O^ A IIET AC UNIOR, Cco^ LEVR WING

3S 5oh E. Tomassia rand HerbertA A.Nelson.



An experimental subeanic flutter investigation of the
effect of root restraint on sweptback, uniform, canti-
lever wange carrying a concentrated weight is re-
ported. The results pertain only to wings wirth
relatively large Length-to-chord ratios where the
restrained root la a small portion of the whole wing.

NACA RM L9KO4

LOW-SPEED INVESTIGATION OF THE AERO-
DYNAM4IC LOADS ON THE DROOP-NOSE FLAP OF
A WING WITH LEADING EDGE SWEPT BACK 47.50
AND HAVING SYMMETRICAL CIRCULAR-ARC AIR-
FOIL SECTIONS AT A REYNOLDS NUMBER OF
4.3 x 106. Edward F. Whittle, Jr. and Marvin P.
Fink. January 6, 1950. 4p. diagra., 3 tbs.
(NACA RMn L9K04)


NACA-langey 1--IS-6.4


F8F-1, SB2C-5, and
F6F-5 airplanes from 1945 to 1949 are analysed
statistically to give the frequency of maximum loads
and airspeeds made in bombing, gunnery, dive pull-
outs, and simulated attacks. Composite V-g dia-
grams are shown for the F4U-4 and TBM-3 air-
planes.


NACA-LEgley 1-4-14 .4M








UNIVERSITY OF FLORIDA


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REPORT xmlns http:www.fcla.edudlsmddaitss xmlns:xsi http:www.w3.org2001XMLSchema-instance xsi:schemaLocation http:www.fcla.edudlsmddaitssdaitssReport.xsd
INGEST IEID EFWRRGJD6_8239E8 INGEST_TIME 2012-03-02T21:00:02Z PACKAGE AA00009235_00017
AGREEMENT_INFO ACCOUNT UF PROJECT UFDC
FILES