Research abstracts

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

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

Notes

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

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 001469326
notis - AGY1019
oclc - 01471285
lccn - 86657025
issn - 0499-9274
Classification:
lcc - TL501 .U5895
System ID:
AA00009235:00018

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National Advisory Committee for Aeronautics



Research Abstracts
NO.57 JANUARY 29, 1954


CURRENT NACA REPORTS
NACA TN 3050

A PHOTOGRAPHIC METHOD FOR DETERMINING
VERTICAL VELOCITIES OF AIRCRAFT IMMEDI-
ATELY PRIOR TO LANDING. Emanuel Rind.
January 1954. 23p. diagrs., photos. (NACA
TN 3050)

A photographic method which has been successfully
used for obtaining statistical data on vertical veloc-
ities of land-based aircraft immediately prior to t
landing contact is described. A long-focal-len '
(40-inch) lens is used. No instrument install
on the aircraft or interference with airport op n
is required. A relatively simple data reducti
/ employed.
I-.\
NACA TN 3059 -5-

ELASTIC BUCKLING UNDER COMBINED STRESS
OF FLAT PLATES WITH INTEGRAL WAFFLE-
LIKE STIFFENING. Norris F. Dow, L. Ross Levin
and John L. Troutman. January 1954. 19p.
diagrs., photos., tab. (NACA TN 3059)

Theory and experiment were compared and found in
good agreement for the elastic buckling under com-
bined stresses of long flat plates with integral
waffle-like stiffening in a variety of configurations.
For such flat plates, 450 waffle stiffening was found
to be the most effective of the configurations for the
proportions considered over the widest range of com-
binations of compression and shear.

NACA TN 3064

DATA ON THE COMPRESSIVE STRENGTH OF
SKIN-STRINGER PANELS OF VARIOUS MATERI-
ALS. Norris F. Dow, William A. Hickman and
B. Walter Rosen. January 1954. 49p. diagram ,
photo., T tabs. (NACA TN 3064)

Flat skin-stringer compression panels of stainless
steel, mild steel, titanium, copper, four aluminum
alloys, and a magnesium alloy were tested. The
results show the effect of variations in yield stress,
Young's modulus, and both yield stress and Young's
modulus for constant yield strain on the buckling,
and load-shortening characteristics of the panels.


NACA TN 3113

ANALYSIS OF STRAIGHT MULTICELL WINGS ON
CAL-TECH ANALOG COMPUTER. Stanley U.
Benscoter and Richard H. MacNeal, California
Institute of Technology. January 1954. 79p.
diagram 4 tabs. (NACA TN 3113)


Using the Cal-Tech analog computer, structural
analyses have been made for four straight multicell
wings. Wings with aspect ratios of 2 and 4 with
rectangular and biconvex cross sections have been
considered. The wings are supported rigidly along
two lines at the faces of the fuselage. Concentrated
loads are applied at the intersection joints of the
ribs and spars. The effe .-oftiearing strains in
the ribs and spars ar cnudhd. Deflections and all
internal force quantiti s have been recorded as well'
_as vibration modes an frequencies. -

^>IAC TN 3116

.QgERE\AATIONS INVOL G'P R'RSSyR FL0
\otaIO HOMOGENEOUS TUR NULE B
SM ide S Uberoi, John i.epkxC University.
Jan ry 954. 61p. diagrs., 2 tabs. (NACA
TN. I

-t is hown that the correlation of fluctuating static
P Essure (in an incompressible and homogeneous
- turbulence) with any fluctuating quantity in the flow
field can be expressed in terms of the correlation of
the same quantity with two or more components of
the velocity. The correlations of pressure with it-
self and of pressure with two velocity components
are investigated in detail for the case of isotropic
turbulence. A postulated relation between the
fourth-order and second-order correlations Is In-
vestigated. The consequences of this relation are
compared with the measurements of the fourth-
order correlations. The root-mean-square pres-
sure and pressure gradients are computed from
second-order correlation for a range of turbulence
Reynolds numbers. Since the pressure gradient is
related to diffusion of marked particles from a
source, the computed pressure-gradient level is
compared with that calculated from a set of diffusion
measurements. The triple correlation equation and
plausible hypotheses relating higher order correla-
tions with second-order correlation are examined
for the possibility of getting a determinate set of
equations for isotropic turbulence.

NACA TN 3118

DESIGN DATA FOR MULTIPOST-STIFFENED
WINGS IN BENDING. Roger A. Anderson, Aldie E.
Johnson, Jr. and Thomas W. Wilder, III. January
1954. 31ip. diagrs., 3 tabs. (NACA TN 3118)

The results of a computational program are pre-
sented which give numerical values of the stiffnesses
required of the various components of a multipost-
stiffened wing to achieve desired buckling-stress
values under bending loads. Two arrangements of
the posts are considered, upright posts and posts
used as diagonals of a Warren truss. This work ex-
tends and summarizes the calculations presented in
NACA RM L52K10a.


AVAILABLEE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASHINGTON vs, D.C., CITING CODE NUMBER ABOVE EACH TITLE
THE REPORT TITLE AND AUTHOR.
4W./s!oV2^








2

NACA TN 3119

STATIC PROPERTIES AND RESISTANCE CHARAC-
TERISTICS OF A FAMILY OF SEAPLANE HULLS
HAVING VARYING LENGTH-BEAM RATIO. Arthur
W. Carter and David R. Woodward. January 1954.
38p. diagrs., photo., 3 tabs. (NACA TN 3119)

The principal results of the investigation of the static
properties of a family of seaplane hulls are pre-
sented in charts from which draft, trim, and upsett-
ing moment may be obtained for wide ranges of load,
center-of-gravity location, and angle of roll. Charts
are presented for the determination of resistance
and trimming moment for length-beam ratios of 6
and 15.

NACA TN 3120

SPAN LOAD DISTRIBUTIONS RESULTING FROM
CONSTANT VERTICAL ACCELERATION FOR THIN
SWEPTBACK TAPERED WINGS WITH STREAM-
WISE TIPS. SUPERSONIC LEADING AND TRAILING
EDGES. Isabella J. Cole and Kenneth Margolis.
January 1954. 62p. diagrs., 2 tabs. (NACA
TN 3120)

On the basis of the linearized supersonic-flow theory,
equations for the span load distribution resulting
from constant vertical acceleration (i.e., linear
angle-of-attack variation with time) are derived for a
series of thin sweptback tapered wings with stream-
wise tips. The analysis is valid, in general, at Mach
numbers for which the wing leading and trailing edges
are supersonic. Computational results are presented
in the form of generalized design curves which per-
mit fairly rapid estimation of the load distribution for
broad ranges of the parameters aspect ratio, taper
ratio, leading-edge sweepback, and Mach number.

NACA TN 3121

SOME EFFECTS OF ASPECT RATIO AND TAIL
LENGTH ON THE CONTRIBUTION OF A VERTICAL
TAIL TO UNSTEADY LATERAL DAMPING AND
DIRECTIONAL STABILITY OF A MODEL OSCIL-
LATING CONTINUOUSLY IN YAW. Lewis R.
Fisher. January 1954. 49p. diagrs., photos.,
6 tabs. (NACA TN 3121)

A fuselage-vertical-tail combination with tails of
two aspect ratios, each of which was tested at four
fuselage tail lengths, was oscillated In yaw through a
range of reduced-frequency parameter correspond-
ing to the lateral motions of airplanes. The phase
lag of the tail force was measured during oscillation
and converted to damping in yaw to determine the ef-
fects of varying these parameters. A complemen-
tary theoretical analysis based on the finite-span
theory of Blot and Boehnlein Is made on the effects
of tail length and aspect ratio on the unsteady lateral
damping and directional stability.

NACA TN 3122

EXPERIMENTAL INVESTIGATION AT A MACH
NUMBER OF 2.41 OF AVERAGE SKIN-FRICTION
COEFFICIENTS AND VELOCITY PROFILES FOR
LAMINAR AND TURBULENT BOUNDARY LAYERS
AND AN ASSESSMENT OF PROBE EFFECTS.
Robert M. O'Donnell. January 1954. 38p. diagre.,
photos. (NACA TN 3122)


NACA
RESEARCH ABSTRACTS NO. 57

Average skin-friction coefficients for laminare-ad \
turbulent flows were measured on a hollow cylinder -
at a Mach number of 2. 41 and over a Reynolds num-
ber range from 0.06 x 106 to 0.95 x 106 per inch.
Comparisons with various laminar- and turbulent-
boundary-layer theories are presented, together with
an assessment of probe effects.

NACA TN 3124

A METHOD FOR ESTIMATING THE EFFECT OF
TURBULENT VELOCITY FLUCTUATIONS IN THE
BOUNDARY LAYER ON DIFFUSER TOTAL-
PRESSURE-LOSS MEASUREMENTS. Jerome Persh
and Bruce M. Bailey. January 1954. 16p. diagra.
(NACA TN 3124)

A method has been devised for estimating the effect
of turbulent velocity fluctuations on diffuser total-
pressure-loss measurements. In the development of
this method, continuity of flow is stipulated, and it is
assumed that the inlet conditions and the diffuser
dimensions are accurately known, that the flow Is
symmetrical, and that the velocity outside the bound-
ary layer at the downstream measuring stations is
not measurably influenced by turbulent velocity fluc-
tuations. Only the case of a conical diffuser with in-
compressible flow is considered, although the proce-
dure may be easily modified to include compressible-
flow considerations.

NACA TN 3131

ON THE KERNEL FUNCTION OF THE INTEGRAL
EQUATION RELATING THE LIFT AND DOWNWASH
DISTRIBUTIONS OF OSCILLATING FINITE WINGS
IN SUBSONIC FLOW. Charles E. Watkins, Harry L.
Runyan and Donald S. Woolston. January 1954.
44p. (NACA TN 3131)

The kernel function of an integral equation relating
the downwash to the lift distribution of a finite wing
oscillating in subsonic compressible flow is treated.
The kernel is reduced to a form which is amenable to
calculations and the types of singularities are shown.
As a check, it is shown that the kernel for the three-
dimensional case reduces exactly to the known kernel
for the two-dimensional case. In addition, results
for the special cases of Mach number of 0 (incom-
pressible case) and frequency of 0 (steady case) are
given.

NACA TN 3132

FATIGUE TESTS AT STRESSES PRODUCING
FAILURE IN 2 TO 10,000 CYCLES. 248-T3 AND
75S-T6 ALUMINUM-ALLOY SHEET SPECIMENS
WITH A THEORETICAL STRESS-CONCENTRATION
FACTOR OF 4.0 SUBJECTED TO COMPLETELY
REVERSED AXIAL LOAD. Herbert F. Hardrath
and Walter Dig. January 1954. 14p. diagrs.,
photo., 2 tabs. (NACA TN 3132)

Notched specimens made of 248-T3 and 75S-T6
aluminum-alloy sheet material, with theoretical
stress-concentration factors equal to 4.0, were sub-
jected to completely reversed axial loads. Failures
occurred in less than 50 cycles at two-thirds of the
static tensile strength and in as few as 2 cycles when
the applied load was near the static strength of the
specimen. The S-N curves were found to be concave
upward for almost the complete range of fatigue
lives; a reversal of curvature occurred at about 10







NACA
RESEARCH ABSTRACTS NO. 57


cycles of load. The fatigue strengths were equiv-
alent for specimens made of each of the two mate-
rials and tested at stresses below 25 ksi; above that
stress the 75S-T6 specimens had the greater fatigue
strength. Compared on the basis of percent of
ultimate tensile strength, the 24S-T3 specimens
were stronger at all stress levels. Test techniques
and special test apparatus are described.


NACA TN 3133

THE FREE-STREAM BOUNDARIES OF TURBULENT
FLOWS. Stanley Corrsin and Alan L. Kistler,
Johns Hopkins University. January 1954. 109p.
diagrs., photos. (NACA TN 3133)

An experimental and theoretical study has been made
of the instantaneously sharp and irregular front
which separates turbulent fluid from contiguous
"nonturbulent' fluid at a free-stream boundary.
The overall behavior of the front is described sta-
tistically in terms of its wrinkle-amplitude growth
and Its lateral propagation relative to the fluid as
functions of downstream coordinate. It is proposed
and justified that the front actually consists of a very
thin fluid layer in which direct viscous forces trans-
mit mean and fluctuating vorticity to previously non-
turbulent fluid. Outside this laminarr superlayer"
there is presumably a field of irrotational velocity
fluctuations (the "nonturbulent" flow) with constant
mean velocity. Theoretical analysis based on this
physical picture gives results which are in plausible
agreement with experimental results for three
turbulent shear flows.




NACA TN 3136

CREEP BENDING AND BUCKLING OF LINEARLY
VISCOELASTIC COLUMNS. Joseph Kempner,
Polytechnic Institute of Brooklyn. January 1954.
22p. diagrs. (NACA TN 3136)

The general dynamic equation of creep bending of a
beam loaded laterally and axially was derived for a
linearly viscoelastic material whose mechanical
properties can be characterized by four parameters.
The material can exhibit instantaneous and retarded
elasticity as well as pure flow. The equation was
used to determine creep characteristics of a beam
in pure bending and a column with an initially sinus-
oldally deformed axis. The results showed that
creep deflection characteristics of the beam are
identical to the creep strain characteristics of a bar
under simple tension or compression. For com-
pressive end loads less than the Euler load, the
column exhibited creep deflections which increase
continuously with time and approach infinity only as
time approaches infinity.


NACA TN 3137

CREEP BENDING AND BUCKLING OF NONLINE-
ARLY VISCOELASTIC COLUMNS. Joseph Kempner,
Polytechnic Institute of Brooklyn. January 1954.
27p. diagrs., 3 tabs. (NACA TN 3137)


3

Differential equations of bending of an idealized
H-section beam column were derived for a nonline-
arly viscoelastic material whose mechanical prop-
erties are analogous to a model consisting of a
linear spring in series with a nonlinear dashpot
whose strain rate is proportional to a power of the
applied stress. The equations were used to obtain
the creep-bending deflections of a beam in pure
bending and of a column with initial sinusoidal devia-
tion from straightness. The results for the simple
beam showed that the deflections vary linearly with
time. The results for the column, with the assump-
tion that the original shape was-maintained, showed
the existence of a finite critical time at which the
deflections become indefinitely large. The critical
time decreases rapidly with increasing axial com-
pression and column inaccuracy.

NACA TN 3138

CREEP BUCKLING OF COLUMNS. Joseph Kempner
and Sharad A. Patel, Polytechnic Institute of
Brooklyn. January 1954. 24p. diagrs., 2 tabs.
(NACA TN 3138)

Formulas are presented for the determination of the
creep deflection-time characteristics of an initially
curved idealized H-section column. These results
were obtained from closed-form solutions of the
differential equation of bending (derived in NACA
TN 3137) of a beam column whose creep properties
are of a nonlinearly viscoelastic nature. The critical
time (the time required for infinite deflections to
develop) established by these solutions Is tabulated
and plotted for a wide range of the parameters in-
volved.

NACA TN 3139

TIME-DEPENDENT BUCKLING OF A UNIFORMLY
HEATED COLUMN. Nathan Ness, Polytechnic
Institute of Brooklyn. January 1954. 18p. diagrs.
(NACA TN 3139)

A theoretical investigation is presented of the time-
temperature- dependent buckling of a pin-jointed
constant-section column, whose initial curvature is
defined by a half-sine wave when the material is
linearly viscoelastic and is heated uniformly along
the column at a prescribed time rate. It was found
that the deviations from straightness Increase with
time and become indefinitely large when heating re-
duces the Young's modulus of the material to the
value at which the applied load is the Euler load of
the column. When the column is heated very rapidly
this critical time represents its limit of usefulness.
When heating takes place less rapidly the deflections
cause bending stresses exceeding the yield stress
of the material at a time considerably smaller than
the critical time.


NACA TN 3157

METHOD FOR CALCULATION OF COMPRESSIBLE
LAMINAR BOUNDARY LAYER WITH AXIAL PRES-
SURE GRADIENT AND HEAT TRANSFER. Paul A.
Libby and Morris Morduchow, Polytechnic Institute
of Brooklyn. January 1954. 44p. diagrs. (NACA
TN 3157)








4


A rapid and sufficiently accurate method, for most
practical purposes, of determining laminar-
boundary-layer characteristics in flow with a given
free-stream Mach number and given velocity distri-
bution at the edge of the boundary layer Is presented.
The method can be easily applied to flow with zero
pressure gradient for any (constant) Prandtl number
of the order of unity and any given temperature dis-
tribution along the wall. For flow in an axial pres-
sure gradient, the method can be applied for a
Prandtl number of unity and any given uniform wall
temperature.

NACA RM E53J07

MEASUREMENT OF HEAT-TRANSFER AND FRIC-
TION COEFFICIENTS FOR FLOW OF AIR IN NON-
CIRCULAR DUCTS AT HIGH SURFACE TEMPERA-
TURES. Warren H. Lowdermilk, Walter F.
Weiland, Jr. and John N. B. Livingood. January
1954. 26p. diagrs. (NACA RM E53J07)

Measurements of average heat-transfer and friction
coefficients were obtained with air flowing through
electrically heated ducts having square, rectangular
(aspect ratio, 5), and triangular cross sections for
a range of surface temperature from 5400 to 17800 R
and Reynolds number from 1000 to 330,000. The re-
sults indicate that the effect of heat flux on correla-
tions of the average heat-transfer and friction coeffi-
cients is similar to that obtained for circular tubes
in a previous investigation and was nearly eliminated
by evaluating the physical properties and density of
the air at a film temperature halfway between the
average surface and fluid bulk temperatures. With
the Nusselt and Reynolds numbers based on the
hydraulic diameter of the ducts, the data for the
noncircular ducts could be represented by the same
equations obtained in the previous investigation for
circular tubes. Correlation of the average differ-
ence between the surface corner and midwall tem-
peratures for the square duct was in agreement with
predicted values from a previous analysis. How-
ever, for the rectangular and triangular ducts, the
measured corner temperature was greater by ap-
proximately 20 and 35 percent, respectively, than
the values predicted by analysis.


NACA TM 1364

THE PLANE PROBLEM OF THE FLAPPING WING.
(Das ebene problem des schlagenden flugels).
Walter Birnbaum. January 1954. 38p. diagrs., tab.
(NACA TM 1364. Trans. from Zeitschrift fiir
augewandte Mathematik und Mechanik, v.4, no.4,
August 1924, p.277-292).

A theoretical study, based on vortex theory as ap-
plied to the linearized equations of motion, is made
of the air forces on wings of infinite aspect ratio in
incompressible flow. Expressions for forces and
moments associated with steady harmonic oscilla-
tions in vertical translation and pitching of wings
are derived in the form of power series In terms of
a reduced frequency parameter. Use is made of the
derived forces first to treat the problem of propul-
ion due to wing flapping and second to determine
theoretical flutter speeds of some simple spring
mounted configurations.


NACA
RESEARCH ABSTRACTS NO. 57

BRITISH REPORTS> -

N-28052*

Royal Aircraft Establishment (Gt. Brit.)
SOME NOTES ON THE DESIGN AND PERFORM-
ANCE OF A THERMAL WATER CONTENT METER
FOR USE IN CLOUDS. J. Rudman and F. J. Bigg.
September 1953. 33p. diagrs., photos. (RAE
Tech. Note Mech. Eng. 145)

Various methods of measuring the liquid water con-
tent of clouds are briefly reviewed and their probable
applications pointed out. It is concluded that the
best type of indicator for use primarily as a warning
instrument in conditions of icing is one of a thermal
type. An instrument is described which consists of
two small heated cylinders, one exposed and the
other in a sheltered position. Free water caught on
the exposed cylinder produces a temperature differ-
ence, relative to the sheltered one, which is used as
a measure of free water content. A theoretical
analysis of the performance of this type of instru-
ment shows that it is Likely to be satisfactory as a
warning device but only good for accurate measure-
ment over a limited range of water concentrations,
unless a resetting process is used.


N-28053*

Royal Aircraft Establishment (Gt. Brit.)
THE ABSORPTIOMETRIC DETERMINATION OF
MAGNESIUM IN TITANIUM METAL. A. Bacon.
September 1953. 27p. diagrs., 11 tabs. (RAE
Tech. Note Met. 177)

Magnesium is separated in alkaline solutions, using
hydrogen peroxide to prevent precipitation of the
titanium. The magnesia is redissolved in standard
acid and determined on the "Spekker" using
Eriochrome Cyanine as the color reagent. Magne-
sium values are given for several samples of
'Kroll" titanium.




N-28054 *

Royal Aircraft Establishment (Gt. Brit.)
CORROSION, STRESS-CORROSION AND FATIGUE
TESTS ON AN ALUMINIUM ALLOY HAVING A
HIGH YOUNG'S MODULUS. G. Meikle,
C. Braithwaite and M. S. Binning. September 1953.
7p. diagr., 2 tabs. (RAE Tech. Note Met. 180)

Corrosion tests have been made on an aluminum
alloy containing 10.5 percent silicon and other ele-
ments to produce a high "E" value. The corrosion
resistance of the alloy was not good, having lost
about 17 percent of the proof stress and about 30
percent of the U. T. S. in 12 months' exposure to sea
water spray. The stress corrosion resistance of
the alloy was not very good. High fatigue results
showed that improvement was imparted to the solu-
tion treated material but not to be fully heat-treated
alloy by polishing. Reverse bend fatigue tests also
showed an improvement by polishing the solution
treated alloy but no improvement by polishing the
fully heat-treated material.







NACA
RESEARCH ABSTRACTS NO.57

N-28055*

Royal Aircraft Establishment (Gt. Brit.)
THE EFFECT OF HEATING SOME ALUMINIUM
ALLOYS FOR SHORT PERIODS UP TO 3400C.
September 1953. 0lp. diagrs., 3 tabs. (RAE
Tech. Note Met. 181)

Samples of 20 S.W.G. (0.036 in.) sheets of DTD.610,
DTD.546, and DTD.687 were heated at 2000, 2500,
3000, and 3400 C for periods from 5 seconds to 32
minutes. The loss in strength of DTD.610 and
DTD.546 at 2000 C is negligible but becomes appreci-
able as the temperature is increased. DTD.687
loses strength at 2000 C even with very short periods
of heating. When heated at 3400 C, resolution oc-
curs In DTD.687 which, on subsequent natural age-
ing, offsets the loss of strength due to softening at
340 C.

N-28057 *

Aeroplane and Armament Experimental Establish-
ment (Gt. Brit.) THE PERFORMANCE OF A
MULTI-ENGINE HELICOPTER FOLLOWING FAIL-
URE OF ONE ENGINE DURING TAKE-OFF OR
LANDING. A. L. Oliver. October 30, 1953. 15p.
diagrs., tab. (AAEE/Res/277)

A theoretical analysis is made of the performance of
multiengine helicopters following failure of one en-
gine in take-off and landing from the type of site
proposed for civil operation in built-up areas. The
performance of a twin engine helicopter of similar
class to the Bristol 173 appears to be just adequate
for safe operation from such a site but the nicety of
handling judgment involved in return landings may
make the performance difficult to achieve. A take-
off technique allowing climb-away after engine fail-
ure at any stage is preferable but this is not possible
for the twin engine machine within the space avail-
able. It is possible if the twin engines are replaced
by four of the same effective total power but only if
a turning climb-away is made after engine failure.
A helicopter with sufficient performance for a
straight climb-away can in general hover with one
engine inoperative.

N-28058*

Royal Aircraft Establishment (Gt. Brit.)
A METHOD OF APPROXIMATE NUMERICAL SOLU-
TION OF NON-LINEAR DIFFERENTIAL EQUA-
TIONS OF THE FORM, it + fl(i, x) + f2(x) = 0.
G. S. Green and H. G. Cuming. September 1953.
16p. diagrs., tab. (RAE Tech. Note GW 277)

A method is suggested for deriving numerically the
essential characteristics of the oscillatory motion
whose nonlinear differential equation is of the form
- + f(i, x) + f2(x) = 0. An example of the applica-
tion of the method to a particular case is given.

N-28099*

Aeronautical Research Council (Gt. Brit.)
ACTUATOR DISC APPLIED TO WALL BOUNDARY
LAYERS IN CASCADES. W. R. Hawthorne and
J. H. Horlock. December 22, 1952. 32p. diagrs.
(ARC 15,490; EA 270)


5
Using the actuator disk theory developed by Bragg
and Hawthorne and Pinsley, a solution for the flow
of a perfect fluid with initial velocity gradients
through a moving cascade is found. The boundary-
layer form factor at exit from the cascade is shown
to be dependent upon the entry boundary-layer
parameters, the exit air angle relative to the cas-
cade, and the ratio of cascade blade speed to the
main stream axial velocity at entry. A comparison
is made between the values of these variables at
which the cascade would stall, and those at which
the flow might be expected to break away from the
wall. The investigation is extended to consider a
complete compressor stage.



N-28100'

Aeronautical Research Council (Gt. Brit.)
SOME ACTUATOR DISC THEORIES FOR THE FLOW
OF AIR THROUGH AN AXIAL TURBO MACHINE.
J. H. Horlock. December 22, 1952. 47p. diagrs.,
tab. (ARC 15,491; EA 271)

Using actuator disk theory, simplified methods are
given for the solution of the direct problem of the
incompressible flow of air through an axial flow
turbomachine. Calculations based on these methods
are compared with other approximate solutions to
the flow through a model compressor stage.




N-28106*

Aeronautical Research Council (Gt. Brit.)
ON THE SOLUTION OF THE NAVIER-STOKES
EQUATIONS FOR A TYPE OF STEADY RADIALLY-
SYMMETRIC VISCOUS FLOW. S. C. R. Dennis.
March 11, 1953. 13p. diagrs. (ARC 15,723;
FM 1878)

The solution of the Navier-Stokes equations for the
steady motion of a viscous fluid near an infinite ro-
tating disk can be made to depend upon the solution
of a pair of simultaneous nonlinear ordinary differ-
ential equations; the same equations govern the flow
between two coaxial rotating disks. In this paper a
method of solution is developed sufficiently to give
some numerical results in a number of related
cases. The results given are not necessarily unique,
for in some cases the equations have been found to
admit of more than one solution. Further solutions
are not considered here, but it is hoped in a later
paper to describe a method of obtaining them.





MISCELLANEOUS


NACA TN 1983

Errata No. 1 on "LONGITUDINAL FLYING
QUALITIES OF SEVERAL SINGLE-ROTOR HELI-
COPTERS IN FORWARD FLIGHT. F. B.
Gustafson, Kenneth B. Amer, C. R. Haig and
J. P. Reeder. November 1949.







NACA
6 RESEARCH ABSTRACTS NO.57

NACA TN 2590 1

Errata No. 2 on 'CALCULATIONS ON THE FORCES
AND MOMENTS FOR AN OSCILLATING WING-
AILERON COMBINATION IN TWO-DEMENSIONAL
POTENTIAL FLOW AT SONIC SPEED. Herbert C.
Nelson and Julian H. Berman. January 1952.





UNPUBLISHED PAPERS


N-15407'

National Bureau of Standards.
THERMODYNAMIC PROPERTIES OF GASEOUS
DIFLUORODICHLOROMETHANE (FREON-12).
Joseph F. Masi. March 20, 1952. i, 19p. diagrs.,
tabs. (National Bureau of Standards. Rept. 1532)

Heat capacities of gaseous difluorodichloromethane
(Freon-12) were measured with the flow calorimeter
from the boiling point of CF2C12 up to 900 C and at
pressures up to 1.5 atmospheres. Tables of the
heat capacity, enthalpy, entropy, and free energy
function of the ideal gas have been calculated for
temperatures from 2000 to 15000 Kelvin.


N-27652'

National Bureau of Standards.
THE REPRESENTATION OF GAS PROPERTIES IN
TERMS OF MOLECULAR CLUSTERS. Harold W.
Woolley. March 1,1952. nl, 1p. dlagrsas., a
(Natronal Bureau of Standards. Rept. 1491)

This report presents a new and more adequate
method for representing the actual thermodynamic
properties of nonpolar gases, through the modlifia-
tion of the usual selection of force law parameters
for the Lennard-Jones type of interaction potential.
Instructions and generalized tables and charts are
included which, together with existing Lables of
second and third viral coefficients, facilitate the
fitting of experlmental data for all nonpolar gases.
Results are shown for A, N2, CO, CO2, CH4, C2H4-


:4







NACA
RESEARCH ABSTRACTS NO.57 7


DECLASSIFIED NACA REPORTS


THE RECLASSIFICATION CHANGES
LISTED IN THIS ISSUE COMPLETE
THE ACTION REQUIRED BY
EXECUTIVE ORDER NO. 10501 ON
REPORTS INITIALLY ISSUED BY
THE NACA AS RESTRICTED.


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



NACA RM 8C29

A RESTRICTED LIST OF AIRCRAFT MATERIALS
RESEARCH PROJECTS. Sponsored by Government
Agencies. May 4, 1948. 95p. (NACA RM 8C29)


This report contains a selected list of Government
sponsored research projects on related aircraft
materials in effect during the calendar year 1947.
Information is contained on titles, description,
sponsoring and conducting agencies.



NACA RM A7DO10

THE DEVELOPMENT OF JET-ENGINE NACELLES
FOR A HIGH-SPEED BOMBER DESIGN. Robert E.
Dannenberg. August 29, 1947. 37p. diagrs.,
photos., tab. (NACA RM A7DIO)


The results of an experimental investigation made
for the purpose of developing suitable jet-engine
nacelle designs for a high-speed medium bomber
are presented. Two types of nacelles were investi-
gated, the first enclosing two 4000-pound-thrust jet
engines and a 65-inch-diameter landing wheel and
the second enclosing a single 4000-pound-thrust jet
engine. Both types of nacelles were tested in posi-
tions underslung beneath the wing and centrally lo-
cated on the wing. The report summarizes the low-
speed investigation and includes some results from
the high-speed portion.



NACA RM A7106

AN EXPERIMENTAL INVESTIGATION OF NACA
SUBMERGED AIR INLETS ON A 1'5-SCALE MODEL
OF A FIGHTER AIRPLANE. Donald E. Gault.
December 5, 1947. 32p. diagrs.. photos., 4 tabs.
(NACA RM A7106)


The results of an experimental investigation of an
NACA submerged inlet system on a 1, 5-scale model
of a fighter airplane are presented. Duct system
total pressure losses and pressure distributions over
the lip and ramp were obtained. It is shown that the
inlet location investigated is unsatisfactory.


NACA RM A7126

PRELIMINARY RESULTS OF A FLIGHT INVESTI-
GATION TO DETERMINE THE EFFECT OF NEGA-
TIVE FLAP DEFLECTION ON HIGH-SPEED
LONGITUDINAL-CONTROL CHARACTERISTICS.
Maurice D. White, Melvin Sadoff, Lawrence A.
Clousing and George E. Cooper. December 16,1947.
22p. diagrs., photos. (NACA RM A7I26)

Flight tests were conducted on two airplanes having
wings of different airfoil section to determine the ef-
fect of deflecting the landing flaps upward on thehigh-
speed longitudinal-control characteristics. For the
airplane with the NACA 66-series airfoil, decrease
in the diving tendency at high Mach numbers resulted
from a reduction in the variation of wing and tail
angle of attack with Mach number with deflected flaps
For the airplane with the NACA 230-series airfoil, a
change in the pitching moment of the airplane without
the tail apparently offset the effect produced by the
reduction of the angle-of-attack variation so that no
appreciable improvement in diving tendencies
resulted with deflected flaps.


NACA RM A7130

AN EXPERIMENTAL INVESTIGATION OF THE
DESIGN VARIABLES FOR NACA SUBMERGED DUCT
ENTRANCES. Emmet A. Mossman and Lauros M.
Randall. January 8, 1948. 59p. diagrs., 3 tabs.
(NACA RM A7I30)


The effect of configuration modifications on the
characteristics of NACA submerged duct entrances
is presented. These tests were made with the duct
inlet installed in the wall of an 8-inch by 36-inch
wind channel. The modifications investigated were
ramp plan form, ramp angle, width-to-depth ratio,
ramp floor shape, deflectors and boundary-layer
thickness.


NACA RM A7J13

AN ANALYSIS OF THE EFFECTS OF WING ASPECT
RATIO AND TAIL LOCATION ON STATIC LONGI-
TUDINAL STABILITY BELOW THE MACH NUMBER
OF LIFT DIVERGENCE. John S. Azelson and
J. Conrad Crown. January 9, 1948. 14p. diagrs.
(NACA RM A7J13)


An analysis of the influence of wing aspect ratio and
tail location on the effects of compressibility on
static longitudinal stability indicates that the use of
reduced wing aspect ratios or short tail lengths lead
to serious reduction in high-speed stability below
the Mach number of lift divergence.


NACA RM A7J22

HIGH-SPEED AERODYNAMIC CHARACTERISTICS
OF A MODEL TAIL PLANE WITH MODIFIED NACA
65-010 SECTIONS AND 00 AND 450 SWEEPBACK.
Joseph L. Anderson and Andrew Martin. January 12,
1948. 85p. diagrs., photo., 2 tabs. (NACA
RM A7J22)










This report presents high-speed aerodynamic char-
acteristics determined from wind-tunnel tests of a
model tail plane having modified NACA 65-010 sec-
tions and a tapered plan form. Results are shown
for the tail plane in an unswept and a 450 sweptback
condition. For the unswept tall the Mach number of
lift divergence was 0.80 while for the sweptback tail
It was aboae 0.875. The report shows how the sub-
stitution of a 450 sweptback horizontal tall for an
unswept tail affects the longitudinal stability and
control forces of an airplane.

NACA RM A7123

HIGH-SPEED AERODYNAMIC CHARACTERISTICS
OF FOUR THIN NACA 63-SERIES AIRFOILS.
Richard J. ILk. December 31, 1947. 53p. diagrs.,
photo., tab. (NACA RM A7J23. Now issued as
TN 2670)

This report contains the results of high-speed wind-
tunnel tests made of the NACA 63-206, 63-208,
63-210, and 63-212 airfoil sections. The results in-
dicate the effect of thickness on the aerodynamic
characteristics of thin NACA 63-series airfoils. The
data of these tests have been compared to similar
data for corresponding NACA 64-series airfoils to
show the effect of thickness distribution on the high-
speed aerodynamic characteristics of NACA 6-series
airfoils. It was concluded that, although the differ-
ences are small, the aerodynamic characteristics of
thin NACA 64-series airfoils are more favorable for
high-speed applications than those of comparable
NACA 63-series airfoils.

NACA RM A7L02

A SUMMARY AND ANALYSIS OF WIND-TUNNEL
DATA ON THE LIFT AND HINGE-MOMENT CHAR-
ACTERISTICS OF CONTROL SURFACES UP TO A
MACH NUMBER OF 0.90. John A. Axelson.
April 30, 1948. 43p. diagrs., tab. (NACA
RM A7L02)


This report contains an extensive summary of wind-
tunnel data showing the liit and hinge-moment char-
acternstics of control surfaces up to a Mach number
of 0.90. The many factors affecting Cho, Ch6.
CL., and CL6 are discussed, including sweep,
airfoil section, aerodynamic balance, control sur-
face profile, and trailing-edge angle.

NACA RM A8H12

HIGH-SPEED AERODYNAMIC CHARACTERISTICS
OF A LATERAL-CONTROL MODEL. I NACA
0012-64 SECTION WITH 20-PERCENT-CHORD
PLAIN AILERON AND 00 AND 450 SWEEPBACK.
Joseph L. Anderson and Walter J. Krumm.
September 27, 1948. 28p. diagrs., photo., 2 tabs.
(NACA RM A8H12)


Wind-tunnel tests were made to determine the aero-
dynamic and lateral-control characteristics of a
semispan wing having the NACA 0012-64 section.
The characteristics were determined through a
large Mach number range for the wing unswept and
swept back 450 and for a 20-percent-chord, plain,
trailing-edge aileron. Results indicate an aileron
overbalance and effectiveness reversal at high sub-
sonic Mach numbers for the wing unswept but only
an overbalance for the wing swept back 450.


NACA RM AUG19


NACA
RESEARCH ABSTRACTS NO.
N-


COMPARISON OF THE AERODYNAMIC CHARAC-
TERISTICS OF THE NACA 0010 AND 0010-64 AIR-
FOIL SECTIONS AT HIGH SUBSONIC MACH NUM-
BERS. Perry P. Polentz. October 7, 1949. 23p.
diagrs., tab. (NACA RM A9Gl9)


Section lift, drag, and pitching-moment character-
istics of the NACA 0010 and NACA 0010-64 airfoil
sections measured at Mach numbers up to 0.91 and
Reynolds numbers between 1.0 x 106 and 1.9 x 106
are presented. Comparisons are made to determine
the principal effects of varying the chordwise loca-
tion of the maximum thickness.



NACA RM A9K22

EXPLORATORY INVESTIGATION OF THE EFFECT
OF SKEWED PLAIN NOSE FLAPS ON THE LOW-
SPEED CHARACTERISTICS OF A LARGE-SCALE
TRIANGULAR-WING-FUSELAGE MODEL.
Bradford H. Wick and David Graham. January 12,
1950. 12p. diagrs., photo. (NACA RM A9K22)


Presented are lift, drag, and pitching-moment data
obtained from tests of a triangular-wing-fuselage
combination with skewed nose flaps. The Reynolds
numbers of the tests were 12.5 and 14.1 million
(based on the wing mean aerodynamic chord). The
semi plan form of the flaps was such that the flap
chord varied from 0 percent wing chord at the model
center line to 100 percent at about 91-percent wing
semispan. It was concluded that skewed nose flaps
of the investigated plan form are a promising means
of delaying both the leading-edge separation and the
tip stalling that occur on a thin, low-aspect-ratio
triangular wing.



NACA RM A9L27

HIGH-SPEED AERODYNAMIC CHARACTERISTICS
OF A LATERAL-CONTROL MODEL. II MODI-
FIED NACA 0012-64 SECTION WITH A 26.6-
PERCENT-CHORD, PLAIN, TRAILING-EDGE
AILERON; WING UNSWEPT AND SWEPT BACK 450.
Walter J. Krumm and Joseph L. Anderson.
March 15, 1950. 55p. diagrs., photo., 2 tabs.
(NACA RM A9L27)


Wind-tunnel tests were made to determine the aero-
dynamic characteristics of a semispan model wing
having a modified NACA 0012-64 section. The
lateral-control characteristics for a 26.6-percent-
chord, plain, trailing-edge aileron were determined
for a Mach number range from 0.40 through 0.925
with the wing unswept and swept back 450.



NACA RM E6I20

ALTITUDE-WIND-TUNNEL INVESTIGATIONS OF
THRUST AUGMENTATION OF A TURBOJET
ENGINE. I PERFORMANCE WITH TAIL-PIPE
BURNING. W. A. Fleming and R. 0. Dietz.
September 25, 1946. 56p. diagrs., photos.
(NACA RM E6I20)


57
1- -*'. >








NACA
RESEARCH ABSTRACTS NO. 57

Engine thrust and fuel consumption were determined
for a wide range of simulated flight conditions and
tail-pipe fuel flows. The investigation was particu-
larly directed toward evaluation of thrust augmenta-
tion for high-speed and high-altitude flight. The
engine tail pipe was modified for the investigation to
reduce the gas velocity at the inlet of the tail-pipe
combustion chamber. The general trends of the
experimental values were in agreement with values
calculated from theoretical equations.

NACA RM E6K27

INVESTIGATION OF SHOCK DIFFUSERS AT MACH
NUMBER 1.85. I PROJECTING SINGLE-SHOCK
CONES. W. E. Moeckel, J. F. Connors and A. H.
Schroeder. June 17, 1947. 47p. diagrs., photos.
(NACA RM E6K27)


Single-shock cones having angles of 200, 300, 400,
500, 60o, and 700 were tested with curved and
straight diffuser-inlet sections. The variation of
total-pressure recovery with tip projection and out-
let area was investigated for each cone to determine
optimum contraction ratios and shock locations.
The effect of angle of attack was also investigated
for several configurations. Maximum total-pressure
recovery was obtained with the 500 cone using the
straight inlet. The highest total-pressure recov-
eries were obtained with subsonic entrance flow.


NACA RM E6L13

INVESTIGATION OF SHOCK DIFFUSERS AT MACH
NUMBER 1.85. II PROJECTING DOUBLE-
SHOCK CONES. W. E. Moeckel, J. F. Connors
and A. H. Schroeder. June 17, 1947. 41p.
diagrs., photos. (NACA RM E6L13)


The total-pressure recovery (at a Mach number of
1.85) of a shock diffuser having projecting cones de-
signed to produce two oblique shocks ahead of the
diffuser inlet was investigated. Four cones with
different included angles were used. Each cone was
investigated with a straight and with a curved
diffuser-inlet section. The effect of angle of attack
and the distribution of static and total pressures at
the diffuser outlet were also investigated for the
best configurations. The highest total-pressure
recoveries were obtained with subsonic inlet flow.
The results were compared with those obtained with
single-shock cones.


NACA RM E6L27a

CHARACTERISTICS OF A HOT JET DISCHARGED
FROM A JET-PROPULSION ENGINE. William A.
Fleming. December 27, 1946. 20p. diagrs.
(NACA RM E6L27a)


An investigation of a heated jet was conducted to
provide information by which an engine could be so
located in an airplane that no external surface is
overheated by the jet. The temperature and the
velocity on the axis of the jet and the diameter of the
jet are presented nondimensionally as functions of
the axial distance from the jet-nozzle outlet and the
diameter of the jet al the vena contract.


9

NACA RM E7BlIlh

COMPUTED TEMPERATURE DISTRIBUTION AND
COOLING OF SOLID GAS-TURBINE BLADES.
J. George Reuter and Carl Gazley. Jr.
February 11, 1947. 13p. diagrs., tab. (NACA
RM E7BIIh)


Computations were made to determine the effects of
gas temperature, blade-root temperature, blade
thermal conductivity, and net gas-to-metal heat-
transfer coefficient on the temperature distribution
in a typical solid turbine blade.


NACA RM E7C12

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
THRUST AUGMENTATION OF A TURBOJET
ENGINE. II PERFORMANCE WITH WATER IN-
JECTION AT COMPRESSOR INLET. Robert 0.
Dietz and William A. Fleming. May 19, 1947.
33p. diagrs. (NACA RM E7C12)


Engine performance at an engine speed of 7600 rpm
was obtained over a wide range of water-air ratios
at pressure altitudes of 5,000 and 20.000 feet and at
a ram-pressure ratio corresponding to a flight
Mach number of about 0.265. A fixed-area tail-pipe
nozzle 16-3 8 inches in diameter was used for this
investigation. Data are presented to show the effect
of water injection on engine performance. A dis-
cussion of the effect of water injection on the pres-
sure and temperature distribution at the compressor
outlet of the turbojet engine is included.


NACA RM E7C26

THE USE OF PERFORATED INLETS FOR EFFI-
CIENT SUPERSONIC DIFFUSION. John C. Evvard
and John W. Blakey. June 25, 1947. 34p. diagrs..
photo. (NACA RM E7C26. Now issued as
RM E51B10)

The use of wall perforations on supersonic diffusers
to avoid the internal contraction-ratio limitation is
described. Experimental results at a Mach number
of 1.85 on a preliminary model of a perforated dif-
fuser having a geometric internal contraction ratio
of 1.49 (the isentropic value) are presented. A
theoretical discussion of the flow coefficients as well
as the size and the spacing of the perforations is also
included. At angles of attack of 00, 30, and 50, total-
pressure recoveries of 0.931, 0.920, and 0.906,
respectively, were obtained.


NACA RM E7D22

A FUEL-DISTRIBUTION CONTROL FOR
CONTINUOUS-FLOW MANIFOLD INJECTION ON
RECIPROCATING ENGINES. Harold Gold and
David M. Straight. June 6, 1947. 16p. diagrs.,
photos. (NACA RM E7D22)


A fuel-distribution control for continuous-flow mani-
fold injection on reciprocating engines is described.
A method of installation of the control on an engine
is suggested.








10

NACA RM E7EI2

PRELIMINARY INVESTIGATION OF EFFECTS OF
GAMMA RADIATION ON AGE-HARDENING RATE
OF AN ALUMINUM-COPPER ALLOY. J. Howard
Kittel. June 20. 1947. 4p. diagr. (NACA
RM E7E12)


A preliminary investigation was made to determine
the effects of gamma radiation on the age-hardening
rate of an aluminum-copper alloy at temperatures of
320 and 700 F. The gamma radiation from a 100-
milligram radium source appeared to have no signif-
icant effect on the age-hardening rate of the alloy.
A metallographic examination of the test specimens
showed no microstructural changes that could be
attributed to the gamma radiation.


NACA RM E7E13

PRELIMINARY INVESTIGATION OF EFFECTS OF
ALPHA-PARTICLE BOMBARDMENT ON THE
CREEP RATE OF ALUMINUM. J. Howard Kittel.
July 3, 1947. 6p. diagrs. (NACA RM E7EI13)


A preliminary investigation was made to determine
the effects of alpha-particle bombardment on the
creep rate of aluminum wire at 4000 F. The alpha
radiation from an 85-millicurie polonium source ap-
peared to decrease slightly the creep rate of the
aluminum. A metallographic examination of the
creep specimens showed no microstructural changes
that could be attributed to the alpha-panrticle om-o-
bardment.


NACA RM E7F10

ALTIl UDE-WIND-TUNNEL INVESTIGATION OF
THRUST AUGMENTATION OF A TURBOJET
ENGINE. III PERFORMANCE WITH TAIL-PIPE
BURNING IN STANDARD-SIZE TAIL PIPE.
William A. Fleming and Richard L. Golladay.
August II. 1947. 47p. diagrs., photos. (NACA
RM E7F101


Thrust augmentation of a turbojet engine by burning
fuel in the tail pipe nas been investigated in the
NACA Cleveland altitude wind tunnel. The engine
performance was determined at several simulated
flight conditions throughout the range of tail-pipe
fuel flows. A tail-pipe combustion chamber having
the same external dimensions as the standard turbo-
jet engine tail pipe was investigated to determine
whether satisfactory operation could be obtained at
high-speed and high-altitude flight conditions. Two
different flame holders were used.



NACA RM E7F II

COMBUSTION-EFFICIENCY INVESTIGATION OF
SPECIAL FUELS IN SINGLE TUBULAR-TYPE
COMBUSTOR AT SIMULATED ALTITUDE CONDI-
TIONS Ralph T. Dittrich. August 15, 1947. 25p.
diagrs 2 tabs (NACA RM E7F111


NACA
RESEARCH ABSTRACTS NO. 57


Ten special straight-run distillate fuels were-inv-et
tigated. The distillates were obtained from various
crude oils and consisted of hydrocarbon mixtures
with distillation temperatures varying from 930 to
6900 F. Three commercial fuels were also tested.
The operating conditions simulated engine operation
at an altitude of 40,000 feet at engine speeds of
7000 and 10,500 rpm. Under certain operating con-
ditions the flame extended beyond the turbine posi-
tion. The tests showed that as the distillation
temperature of the fractions from the same crude
increased, the combustion efficiency decreased.



NACA RM E7F12

FUEL INVESTIGATION IN A TUBULAR-TYPE
COMBUSTOR OF A TURBOJET ENGINE AT SIMU-
LATED ALTITUDE CONDITIONS. Adelbert 0.
Tischler and Ralph T. Dittrich. August 1, 1947.
41p. diagrs., photo., 2 tabs. (NACA RM E7F12)


A series of 11 fuels, which ranged in volatility from
gasoline to Diesel oil and which included hydro-
carbons of the paraffinic, naphthenic, olefinic, and
aromatic types, was tested in a single tubular com-
bustion chamber of a turbojet engine under inlet-air
conditions that simulated engine operation at two
engine speeds at an altitude of 40,000 feet. Tests
were also conducted at two additional inlet-air con-
ditions. Temperature-rise data at various fuel-air
ratios were obtained for each set of air-flow condi-
tions.



NACA RM E7G03

EXPERIMENTAL INVESTIGATION OF PERFORM-
ANCE AND OPERATING CHARACTERISTICS OF A
TAIL-PIPE BURNER FOR A TURBOJET ENGINE.
David S. Gabriel, E. Vincent Martinson and Robert
H. Essig. October 30, 1947. 29p. diagrs. (NACA
RM E7G03)

Presents description and operating characteristics
of 10 full-scale tail-pipe burners. The combustion
and pressure-drop characteristics of most satisfac-
tory burner were investigated. A tail-pipe burner
was developed that operated satisfactorily over a
range of fuel-air ratios with inlet conditions of gas
temperature and velocity simulating those in a
typical turbojet engine,




NACA RM E7I25a

AN ANALYSIS OF CONTROL REQUIREMENTS AND
CONTROL PARAMETERS FOR DIRECT-COUPLED
TURBOJET ENGINES. David Novik and Edward W.
Otto. February 18, 1948. 50p. diagrs. (NACA
RM E7I25a)

Presents discussion of control parameters for
direct-coupled turbojet engine, based on analysis of
steady-state operation, acceleration, and decelera-
tion, and describes operation of a hypothetical con-
trol system.







NACA
RESEARCH ABSTRACTS NO. 57

NACA RM E7J19

EXPERIMENTAL INVESTIGATION OF THRUST
AUGMENTATION OF 4000-POUND-THRUST
CENTRIFUGAL-FLOW-TYPE TURBOJET ENGINE
BY INJECTION OF WATER AND ALCOHOL AT
COMPRESSOR INLETS. William L. Jones and
Helmuth W. Engelman. May 14, 1948. 23p. diagrs.
(NACA RM E7J19)


Contains curves and data to determine amount of
thrust augmentation obtainable by Injection of vari-
ous water and alcohol mixtures at compressor inlets
of a 4000-pound-thrust turbojet engine at standard
NACA conditions. A maximum thrust augmentation
of 26 percent was obtained by injection of 4.5pounds
per second of water and 2.0 pounds per second of
alcohol.


NACA RM E7K14

EXPERIMENTAL INVESTIGATION OF THRUST
AUGMENTATION OF AXIAL-FLOW-TYPE 4000-
POUND-THRUST TURBOJET ENGINE BY WATER
AND ALCOHOL INJECTION AT COMPRESSOR IN-
LET. Burnett Baron, Harry W. Dowman and
William C. Dackis. July 8, 1948. 41p. diagrs.,
photos. (NACA RM E7K14)


Reports on experimental investigation of thrust aug-
mentation of axial-flow-type turbojet engine with a
4000-pound-thrust rating by means of water-alcohol
Injection at the compressor inlet. Augmentation was
limited by centrifugal separation of liquid and air in
the compressor and was sensitive to engine inlet-air
temperature. A thrust augmentation of 15.4 percent
was obtained with injection of 3 pounds per secondof
water alone at an inlet-air temperature of 5480 F.


NACA RM E7L12

RELATION OF NOZZLE-BLADE AND TURBINE-
BUCKET TEMPERATURES TO GAS TEMPERA-
TURES IN A TURBOJET ENGINE. J. Elmo Farmer.
April 30, 1948. 37p. diagrs., photos. (NACA
RM E7L12)

Presents results of investigation to determine exper-
imentally turbine-nozzle-blade and turbine-bucket
temperatures in a turbojet engine and to correlate
these temperatures with gas temperatures. Maxi-
mum indicated temperatures were about 19000 F for
the nozzle blade and 15000 F for the turbine bucket.
The maximum turbine-nozzle-blade temperature was
800 to 2700 F higher than the calculated average
turbine-inlet-gas temperature; the maximum turbine-
bucket temperature was about 1500 F less than the
calculated average turbine-inlet-gas temperature.


NACA RM E7L16

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
THRUST AUGMENTATION OF A TURBOJET
ENGINE. IV PERFORMANCE WITH TAIL-PIPE
BURNING AND WATER INJECTION. Robert 0.
Diets, Jr., George Wishnek and John K.L Kuenzsig.
June 15, 1948. 31p. diagrs., photo. (NACA
RM E716)


II

Results of investigation of thrust augmentation of
axial-flow-type turbojet engine by combination of
tail-pipe burning and water injection are presented.
Thrust increases obtainable with engine-inlet water
Injection reached a peak at a water-air ratio of 0.035
and decreased with larger amounts of injection;
whereas thrust Increases obtainable with combustion-
chamber water injection increased as water-air
ratio was raised throughout range of water-air ratios
investigated. Greater maximum thrust increases
were available with combustion-chamber water
injection than with engine-inlet water injection.
Thrust increases were accompanied by large in-
creases in specific liquid consumption.


NACA RM E7Ll7

CONTROL DURING STARTING OF GAS-TURBINE
ENGINES. Robert J. Koemng and Marcel Dandois.
June 17, 1948. 39p. diagrs., photos. (NACA
RM E7L17)

Reports on investigation of variables pertinent to the
control of gas temperatures during starting of gas-
turbine engines. Results indicate that poor control
of gas temperatures during starting is caused by an
accumulation of fuel in the engine before ignition and
by excessive fuel-flow rates at the time of ignition.
Prompt ignition was obtained by use of well atomized
fuel sprays.



NACA RM E7Ll8

VIBRATION OF TURBINE BLADES IN A TURBOJET
ENGINE DURING OPERATION. W. C. Morgan, R. H.
Kemp and S. S. Manson. April 22, 1948. 17p.
diagrs.. photos. (NACA RM E7Lb1)


An experimental investigation was made to determine
the vibration phenomena that occur in gas-turbine
blades during service operation. Through the use of
high-temperature strain gages, vibratory phenomena
in turbine blades were observed and evaluated in
terms of modes, frequencies, and stress range. The
frequencies of the principal vibrations were found to
be related to the number of nozzle blades and com-
bustion chambers.




NACA RM E7L30

COMBUSTION-EFFICIENCY AND ALTITUDE-LIMIT
INVESTIGATIONS OF FIVE FUELS IN AN ANNULAR
TURBOJET COMBUSTOR. Jerrold D. Wear and
Edmund R. Jonash. June 7, 1948. 19p. diagrs.
(NACA RM E7L30)

Five fuels of various boiling temperatures and vari-
ous hydrocarbon types were Investigated in a jet-
propulsion annular combustor of 10-3 8 inch diam-
eter to determine the effect of boiling temperature
and paraffinic and aromatic hydrocarbon types on
combustion efficiency and altitude operational limit
at three inlet-air conditions. The performance dif-
ference among the fuels was greatest at inlet-air
conditions characterized by unstable combustion.








12
NACA RM E8A28a

BENCH AND ENGINE OPERATION OF A FUEL-
DISTRIBUTION CONTROL. Harold Gold and Robert
J. Koenig. June 14, 1948. 34p. diagrs., photos.
(NACA RM E8A28a)


Presents application of a fuel-distribution-control
method to actual gas-turbine engine operation. The
control used was designed to equalize the flow to the
14 nozzles of a gas-turbine engine. The maximum
measured deviation from perfect distribution during
engine operation was 3.8 percent. It was shown that
the control model is capable of maintaining this
accuracy independently of changes in fuel-nozzle
resistances from 0 up to 1.46 times the resistance of
a normal engine fuel nozzle.

NACA RM E8CO2a

EFFECT OF FUELS AND FUEL-NOZZLE CHARAC-
TERISTICS ON PERFORMANCE OF AN ANNULAR
COMBUSTOR AT SIMULATED ALTITUDE CONDI-
TIONS. Richard J. McCafferty. September 28,
1948. 90p. diagrs., photos., tab. (NACA RM E8C02a)


The effect of fuels and fuel-nozzle characteristics on
performance of an annular combustor at simulated
altitude conditions was determined by operating with
different fuels and fuel-nozzle sizes with the com-
bustor inlet-air conditions each independently altered
near the altitude operational limits. Results indicate
that amount of difference in performance at altitude
among different fuels is determined by hydrocarbon
type and fuel boiling points. Reduction in fuel-nozzle
size increased combustion efficiency at low heat-
input values but produced lower temperature-rise
limits.


NACA RM E8C08

A FUEL-DISTRIBUTION CONTROL FOR GAS-
TURBINE ENGINES. Harold Gold and David M.
Straight. June 14, 1948. 17p. diagrs., photos.
(NACA RM E8C08)


A device to control the distribution of fuel to any
number of discharge nozzles of a gas-turbine engine
is presented. A model of the device controlled the
flow to four discharge nozzles within 2 percent of
perfect distribution over a wide range of fuel flows
and was unaffected by uneven discharge-nozzle pres-
sures.


NACA RM E8D12

INVESTIGATION OF HIGH-TEMPERATURE OPERA-
TION OF LIQUID-COOLED GAS TURBINES. I -
TURBINE WHEEL OF ALUMINUM ALLOY, A HIGH-
CONDUCTIVITY NONSTRATEGIC MATERIAL.
Harry Kottas and Bob W. Sheflin. July 22, 1948.
15p. diagrs., photos., tab. (NACA RM E8D12)


A liquid-cooled gas turbine with aluminum-alloy
blades and disk was built to investigate the effect of
liquid cooling and high-conductivity metals on limit-
ing turbine-Inlet gas temperature. The turbine was
successfully operated at a speed of 19,000 rpm and an
inlet gas temperature of 21000 F.


NACA
RESEARCH ABSTRACTS NO.57

NACA RM E8DI4 "
~ -'I-*
GAS-TURBINE-ENGINE OPERATION WITH
VARIABLE-AREA FUEL NOZZLE. Harold Gold and
David M. Straight. July 9, 1948. 47p. diagrs.,
photos. (NACA RM E8D14)


The characteristics of variable-area and fixed-area
atomizing nozzles are discussed in relation to use in
aircraft gas-turbine engines. A variable-area
nozzle and a fuel-distribution control that were used
in the operation of a gas-turbine engine are de-
scribed. When the engine was equipped with variable
area fuel nozzles, thrust specific fuel consumption in
the lower half of the fuel-flow range was reduced,
fuel pressures were lowered in the upper half of the
fuel-flow range, and starting characteristics were
improved.


NACA RM E8D23

PERFORMANCE OF SEVERAL AIR EJECTORS
WITH CONICAL MIXING SECTIONS AND SMALL
SECONDARY FLOW RATES. S. C. Huddleston, H. D.
Wilsted and C. W. Ellis. July 19, 1948. 74p. diagrs.
(NACA RM E8D23)


Experimental results presented are limited to inves-
tigations of conical-type mixing-section ejectors at
ratios of mixing-section minimum diameter to
primary-jet-nozzle diameter of 1.21, 1.10, and 1.00
using unheated air. Results were cross-plotted in
charts to show the performance of all ejectors within
the range of configurations investigated. For diam-
eter ratios of 1.21 and 1.10, the spacing giving max-
imum airflow varied with diameter ratio but did not
vary with primary and secondary pressure ratios.
In general, the ejector had little effect on the primary
nozzle thrust.


NACA RMI E8D29

EFFECT OF MACH NUMBER ON PERFORMANCE
OF AN AXIAL-FLOW COMPRESSOR ROTOR-BLADE
ROW. Paul D. Dugan, John J. Mahoney and William
A. Benser. September 28, 1948. 27p. diagrs., photo.
(NACA RM E8D29)


An investigation was conducted to study the influence
of high relative Inlet Mach number on the perform-
ance of a highly loaded rotor blade row. A range of
weight flow was investigated for equivalent tip
speeds ranging from 300 to 915 feet per second,
which correspond to relative inlet Mach numbers of
0.30 to 0.90 at the mean radius. At high Mach num-
bers, above 0.70, the blade-section-performance re-
sults of this investigation indicated good correlation
with low-speed cascade results. The efficiency
gradually decreased from 0.950 to 0.915 as the rela-
Live inlet Mach number increased from 0.30 to 0.82.


NACA RM E8E10

ANALYTICAL INVESTIGATION OF EFFECT OF
WATER-COOLED TURBINE BLADES ON PERFORM-
ANCE OF TURBINE-PROPELLER POWER PLANTS.
William D. Bowman. August 16, 1948. 67p. diagrs.
(NACA RM E8E10)







NACA
RESEARCH ABSTRACTS NO.57

Performance of six turbine-propeller power plants,
each having a multistage turbine equipped with water-
cooled turbine blades and designed for maximum tur-
bine Inlet temperatures of 20000, 25000, 30000, 40000,
and 45000 R, respectively, is calculated over a range
of flight speeds from 200 to 600 miles per hour at
altitudes from sea level to 35,000 feet. High temper-
ature operation possible with water-cooled blades
yields, at appropriate pressure ratios, performance
gains that are not vitiated by losses incurred in cool-
ing process. The heat-exchanger capacity and size
are approximately 20 percent of that of a liquid-
cooled piston engine of equal rated power.


NACA RM E8F04

DESIGN AND PERFORMANCE OF EXPERIMENTAL
AXIAL-DISCHARGE MIXED-FLOW COMPRESSOR.
I IMPELLER DESIGN THEORY. Arthur W.
Goldstein. August 12, 1948. 32p. diagrs. (NACA
RM E8F04)

An axial-discharge centrifugal compressor, especial-
ly adapted for jet engines because of the large mass
flow per unit frontal area, is described. General
equations of relative Qfluid motion are developed to
show assumptions involved and empirical character
of simplifications. The best impeller was selected
on basis of maximum air-flow capacity, which was
19.6 pounds per second for 14-inch diameter with
tip speed of 1480 feet per second and pressure ratio
of 3.5.


NACA RM E8FO9

ALTITUDE-WIND-TUNNEL INVESTIGATION OF A
4000-POUND-THRUST AXIAL-FLOW TURBOJET
ENGINE. I PERFORMANCE AND WINDMILLING
DRAG CHARACTERISTICS. William A. Fleming.
August 3, 1948. 76p. diagrs., photos. (NACA
RM E8F09)

Performance of a 4000-pound-thrust axial-flow
turbojet engine was investigated at altitudes from
5000 to 40,000 feet and at ram pressure ratios from
1.02 to 1.86. The specific fuel consumption based
on net thrust horsepower decreased rapidly with air-
speed. Windmilling drag of the engine was relative-
ly high and the inlet should be closed when the engine
is inoperative in flight. The results show that an
accurate calculation of the jet thrust of the engine
can be made from measurements of the tempera-
tures and pressures obtained from a survey across
the jet-nozzle exit.


NACA RM E8FO9a

ALTITUDE-WIND-TUNNEL INVESTIGATION OF A
4000-POUND-THRUST AXIAL-FLOW TURBOJET
ENGINE. II OPERATIONAL CHARACTERISTICS.
William A. Fleming. August 6, 1948. 53p. diagrs.,
photos., 3 tabs. (NACA RM E8FO9a)

Operational characteristics of a 4000-pound-thrust
axial-flow turbojet engine were obtained at pressure
altitudes from 5000 to 50,000 feet, ram pressure
ratios from 1.00 to 1.86, and temperatures from 600
to -500 F. The effects of altitude and airspeed on
such operational characteristics as combustion sta-
bility acceleration, starting, and fuel-control system
were studied. The engine could be operated at full


13

speed without serious burner unbalance at altitudes
up to 50,000 feet. Acceleration of the engine was
relatively slow and the time required for accelera-
tion increased with altitude. The engine started
normally and easily below 20,000 feet.




NACA RM E8FO9b

ALTITUDE-WIND-TUNNEL INVESTIGATION OF A
4000-POUND-THRUST AXIAL-FLOW TURBOJET
ENGINE. M PERFORMANCE CHARACTERISTICS
WITH THE HIGH-FLOW COMPRESSOR. William A.
Fleming and Richard L. Golladay. August 5, 1948.
69p. diagrs., photos. (NACA RM E8FO9b)


An investigation was conducted to determine the per-
formance of a 4000-pound-thrust axial-flow turbojet
engine with a high-flow compressor at altitudes from
5000 to 40,000 feet and ram pressure ratios from
1.00 to 1.82. Engine performance with high-flow and
low-flow (standard) compressors is compared. In-
stallation of a high-flow compressor in the engine in
place of a low-flow compressor at a corrected engine
speed of 7600 rpm and a ram pressure ratio of 1.40
gave an increase in corrected jet thrust of 11 per-
cent, in corrected net thrust of 6.5 percent, in
corrected fuel consumption of 14 percent, and in
corrected airflow of 12 percent.





NACA RM E8F09c

ALTITUDE-WIND-TUNNEL INVESTIGATION OF A
4000-POUND-THRUST AXIAL-FLOW TURBOJET
ENGINE. IV ANALYSIS OF COMPRESSOR PER-
FORMANCE. Robert 0. Dietz, Jr. and Frank L.
Suozzi. August 5, 1948. 54p. diagrs., photos., tab.
(NACA RM E8FO9c)


Performance of two 11-stage axial-flow compres-
sors operating in a 4000-pound-thrust axial-flow
turbojet engine was investigated at simulated alti-
tudes from 5000 to 40,000 feet, ram pressure ratios
from 1.00 to 1.86, and compressor Mach numbers
from 0.24 to 0.95. One was a standard compressor
and the other, a similar compressor except that the
blade angles of the rotor and stator blades were in-
creased about 5o to obtain greater airflow. Max-
imum adiabatic efficiency for both compressors was
about 85 percent. Increasing the blade angles 50
increased the airflow approximately 10 percent.
Efficiency and pressure ratio al a given compressor
Mach number were unaffected by changes in
Reynolds number.





NACA RM E8FO9d

ALTITUDE-WIND-TUNNEL INVESTIGATION OF A
4000-POUND-THRUST AXIAL-FLOW TURBOJET
ENGINE. V ANALYSIS OF TURBINE PERFORM-
ANCE. Richard P. Krebs and Reece V. Hensley.
August 4, 1948. 27p. diagrs., photo. (NACA
RM E8FO9d)








14

Performance characteristics of a turbine from a
4000-pound-thrust axial-flow turbojet engine were
obtained at altitudes from 5000 to 40,000 feet and at
ram pressure ratios from 1.01 to 1.77. Turbine
efficiency was practically unaffected by changes in
altitude or ram pressure ratio. The mammum
efficiency obtained was about 82 percent.


NACA RM E8F17

PERFORMANCE INVESTIGATION OF CAN-TYPE
COMBUSTOR. I INSTRUMENTATION, ALTITUDE
OPERATIONAL LIMITS AND COMBUSTION EFFI-
CIENCY. Eugene V. Zettle and William P. Cook.
September 16, 1948. 21p. diagrs., photo., tab.
(NACA RM EBF17)


Investigation of single can-type combustor of turbo-
jet engine was conducted to determine altitude oper-
ational limits with two fuels, combustion efficiencies
at various altitudes and engine speeds, combustor-
outlet temperature distribution, and combustor total-
pressure drop. Limits with AN-F-32 fuel were ap-
proximately 60,000 feet for engine speed of 6000 rpm
and 38,000 feet for engine speed of 4000 rpm. Alti-
tude operational limits with AN-F-32 fuel are higher
over largest part of engine-speed range than with
AN-F-28 fuel.


NACA RM E8F29

OPERATING TEMPERATURES OF 1-40-5 TURBO-
JET ENGINE BURNER LINERS AND THE EFFECT
OF TEMPERATURE VARIATION ON BURNER-
LINER SERVICE LIFE. H. D. Wilsted, Robert T.
Duffy and Ralph E. Grey. August 23, 1948. 25p.
diagrs., photos. (NACA RM E8F29)

The metal temperature and the temperature distrib-
ution in an 1-40-5 turbojet engine were determined
over the operating range of engine speeds using ther-
mocouples welded to the outer surface of the burner
liner. A ceramic-coated thermocouple was used
throughout the investigation. Burner-liner temper-
atures increased at an increasing rate with engine
speed; the highest temperature recorded was 15000 F.
Temperature gradients were as large as 7000 F per
inch. The high local temperature gradients were ap-
parently the cause of short burner-liner service life.


NACA RM E8G02

INVESTIGATION OF THE 1-40 JET-PROPULSION
ENGINE IN THE CLEVELAND ALTITUDE WIND
TUNNEL. I PERFORMANCE AND WINDMILLING
DRAG CHARACTERISTICS. Stanley L. Gender and
William K. Koffel. August 24, 1948. 75p. diagrs.,
photos. (NACA RM E8G02)


An altitude-wind-tunnel investigation of the perform-
ance and windmilling characteristics of an 1-40 jet-
propulsion engine was conducted over a range of sim-
ulated flight conditions to determine the effects of
altitude and ram pressure ratio. The use of general-
izing factors for estimating altitude performance of
the engine gave only fair results. Specific fuel con-
sumption based on net thrust horsepower showed no
effect of altitude but decreased with Increasing ram
pressure ratio at an engine speed of 11,500 rpm.


NACA
RESEARCH ABSTRACTS NO. 57

NACA RM E8GO2a s.
-a'. *.
INVESTIGATION OF THE 1-40 JET-PROPULSION
ENGINE IN THE CLEVELAND ALTITUDE WIND
TUNNEL. II ANALYSIS OF COMPRESSOR PER-
FORMANCE CHARACTERISTICS. Robert 0. Dietz,
Jr. and Robert M. Geisenheyner. August 26, 1948.
28p. diagrs., photos. (NACA RM E8G02a)


Performance characteristics of a centrifugal com-
pressor operating as an integral part of a jet propul-
sion engine were obtained for a range of compressor
Mach numbers from 0.72 to 1.46, ram pressure
ratios from 0.98 to 1.76, and simulated altitudes from
10,000 to 40,000 feet. From results obtained over a
wide range of altitudes, it was determined that the
compressor performance is primarily dependent on
the compressor-inlet Mach number. Variations of
Reynolds number of the air at the compressor inlet
had little effect on compressor performance.


NACA RM E8G002b

INVESTIGATION OF THE 1-40 JET-PROPUSION
ENGINE IN THE CLEVELAND ALTITUDE WIND
TUNNEL. I ANALYSIS OF TURBINE PERFORM-
ANCE AND EFFECT OF TAIL-PIPE DESIGN ON
ENGINE PERFORMANCE. Richard P. Krebs and
Frederick C. Foshag. August 26,1948. 27p. diagrs.,
photo. (NACA RM E8GO2b)


Performance characteristics of a turbine operating
as an integral part of a jet-propulsion engine were
determined for a range of simulated altitudes from
10,000 to 40,000 feet and ram pressure ratios from
0.98 to 1.76. A comparison of engine performance
with three different tall pipes is also presented.
Turbine efficiency was unaffected by changes in
altitude or the tail-pipe designs but varied with
changes in ram pressure ratio. The most satisfac-
tory engine performance was obtainedwith a constant-
diameter tail pipe having a short nozzle at the outlet.


NACA RM E8G02c

INVESTIGATION OF THE 1-40 JET-PROPULSION
ENGINE IN THE CLEVELAND ALTITUDE WIND
TUNNEL. IV ANALYSIS OF COMBUSTION-
CHAMBER PERFORMANCE. Reece V. Hensley.
August 25, 1948. 34p. diagrs. (NACA RM E8GO2c)


Combustion-chamber performance characteristics
have been determined from an 1-40 jet-propulsion
engine over a range of simulated altitudes and ram
pressure ratios. Combustion efficiency varied
directly with ram pressure ratio and engine speed
and varied inversely with altitude. Percentage
losses in total pressure were appreciably affected
by changes in altitude or ram pressure ratio up to a
value of 1.3.


NACA RM E8GO2d

INVESTIGATION OF THE 1-40 JET-PHOPULSION
ENGINE IN THE CLEVELAND ALTITUDE WIND
TUNNEL. V OPERATIONAL CHARACTERISTICS.
Richard L. Golladay and Stanley L. Gender.
August 25, 1948. 71p. diagrs., photos., 6 tabs.
(NACA RM E8GO2d)







NACA
RESEARCH ABSTRACTS NO. 57

Operational characteristics of a jet-propulsion en-
gine determined at pressure altitudes from 10,000 to
50,000 feet and ram pressure ratios from 1.00 to
1.76 include effects of altitude and airspeed on oper-
ating speed range, starting, windmilling, accelera-
tion, speed regulation, cooling, and vibration of the
standard and modified engines. Maximum engine
speed was obtainable at all altitudes and airspeeds
with each fuel-control system investigated. The
maximum idling speed was raised by increases in
altitude and airspeed. Ignition at 30.000 feet was
difficult. Windmilling starts were not successful.
The engine speed was held constant during simulated
dives and climbs at constant throttle position.


NACA RM E8H11

THEORETICAL COMPARISON OF SEVERAL
METHODS OF THRUST AUGMENTATION FOR
TURBOJET ENGINES. Eldon W. Hall and E. Clinton
Wilcox. October 27, 1948. 40p. diagrs. (NACA
RM E8HII. Now issued as Rept. 992)


Performance of several methods of thrust augmenta-
tion for turbojet engines was calculated for flight
Mach numbers from 0 to 2.5 and for altitudes of sea
level and 35,332 feet. The methods investigated were
tail-pipe burning, water injection at compressor inlet,
combination of tail-pipe burning and water injection,
bleedoff, and rocket assist. Results indicated that
tail-pipe burning plus water injection was best for
large amounts of thrust augmentation and tail-pipe
burning was best for smaller amounts inasmuch as
these methods have the lowest ratio of augmented-to-
normal total liquid consumption for a given thrust
increase.

NACA RM E8H12

DETERMINATION OF AVERAGE HEAT-TRANSFER
COEFFICIENTS FOR A CASCADE OF SYMMETRI-
CAL IMPULSE TURBINE BLADES. I HEAT
TRANSFER FROM BLADES TO COLD AIR. Gene L.
Meyer. 41p. diagrs., photos. (NACA RM E8HI2)


A cascade of symmetrical impulse turbine blades
was investigated to determine average surface heat-
transfer coefficients. The results were correlated
with other cascade data [or impulse blades and are
presented in dimensionless form. Heat-transfer
coefficients based on an effective gas temperature
were independent of Mach number.


NACA RM E8103

EFFECT OF PRESSURE RATIO AND INLET PRES-
SURE ON PERFORMANCE OF EXPERIMENTAL
GAS TURBINE AT INLET TEMPERATURE OF
8000 R. Robert C. Kohl and Robert G. Larkin.
November 22, 1948. 7p. diagrs. (NACA RM E8103)


An experimental gas turbine was operated over a
range of blade-jet speed ratios, total pressure ratios,
and inlet total pressures at a constant inlet temper-
ature of 8000 R. Peak over-all efficiencies were
obtained at blade-jet speed ratios from 0.525 to
0.575 for all runs. The variation in peak efficiency
with inlet pressure and pressure ratio was of small
magnitude for the conditions investigated.


15
NACA RM E8121

SOME EFFECTS OF STATOR CONE ANGLE AND
BLADE-TIP LEAKAGE ON 40-PERCENT-
REACTION TURBINE HAVING ROTOR-BLADE
CAPS. Robert E. English, Robert J. McCready and
John S. McCarthy. March 23, 1949. 28p. diagrs.,
photo. (NACA RM E8I21)


A turbine having 40-percent reaction and rotor-blade
caps, which form a continuous rotating cylindrical
shroud, was operated with two stators having cone
angles of 700 and 00 and with two stationary shrouds-
a labyrinth, no-leakage shroud and a cylindrical
shroud. With the 00-cone-angle stator, the efficien-
cy was 0.04 higher than with the 700-cone-angle
stator. Replacing the labyrinth shroud with the
cylindrical shroud having a radial clearance 0.016 of
the blade height produced no measurable change in
efficiency.



NACA RM E8J22

VIBRATION SURVEY OF BLADES IN 10-STAGE
AXIAL-FLOW COMPRESSOR. I STATIC INVES-
TIGATION. Andre J. Meyer, Jr. and Howard F.
Calvert. January 31, 1949. 34D. diagrs., photos.,
2 tabs. (NACA RM E8J22)


Natural frequencies were measured in the blading of
a 10-stage axial-flow compressor and these frequen-
cies were compared with possible exciting forces.
Node shapes of higher modes of vibration were deter-
mined from sand patterns in an attempt to correlate
the position of high-stress points for the various
modes with the location of actual failures in the
seventh- and tenth-stage blades. Critical-speed
diagrams were plotted to show possible cause of
failures.



NACA RM E8J22a

VIBRATION SURVEY OF BLADES IN 10-STAGE
AXIAL-FLOW COMPRESSOR. II DYNAMIC IN-
VESTIGATION. Andre J. Meyer, Jr. and Howard F.
Calvert. January 31, 1949. 25p. diagrs., photos.
(NACA RM E8J22a)


Vibratory stresses were measured in the first five
stages by the use of strain gages mounted on the
blades of a 10-stage axial-flow compressor. Effects
of surge and of changes in pressure ratio on ampli-
tude of blade vibrations were investigated. Goodcor-
relation was obtained between first bending-mode
frequencies that were dynamically measured and
those that were calculated by correcting static
measurements to account for the effects of centrif-
ugal force.



NACA RM E8J22b

VIBRATION SURVEY OF BLADES IN 10-STAGE
AXIAL-FLOW COMPRESSOR. III PRELIMINARY
ENGINE INVESTIGATION. Andre J. Meyer, Jr. and
Howard F. Calvert. January 31, 1949. 23p.diagrs.,
photo. (NACA RM E8J22b)








16

Vibratory stresses were measured in the blading of
a 10-stage axial-flow compressor under operating
conditions. Curves are presented showing maximum
allowable vibratory stress for a given speed, change
of damping coefficient with mounting of a strain gage
at base of blade, effect of rotor speed on bladenatu-
ral frequency, and effect of order of the first bending-
mode vibration on stress. For all stages, the lower
the order of vibration the higher the stress but no
destructive vibrations were detected.

NACA RM E8J25

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
OPERATIONAL CHARACTERISTICS OF
WESTINGHOUSE X24C-4B AXIAL-FLOW TURBO-
JET ENGINE. W. Kent Hawkins and Carl L. Meyer.
November 23, 1948. 39p. diagrs., photos. (NACA
RM E8J25)
Operating characteristics of Westinghouse X24C-4B
axial-flow turbojet engine presented include: Oper-
able engine-speed range at altitude, acceleration and
deceleration, altitude and flight-Mach-number com-
pensation of governor, lubrication-system perform-
ance, and starting characteristics. Operable range
of engine speeds was limited at high altitudes and
low flight Mach numbers. Accelerations and decel-
erations at altitudes above 25,000 feet were limited
by combustion blow-out. Altitude and flight-Mach-
number compensation of governor was good at high
engine speeds and starting characteristics of one
engine configuration were very unsatisfactory.


NACA RM E8J25a

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
PERFORMANCE AND WINDMILLING DRAG CHAR-
ACTERISTICS OF WESTINGHOUSE X24C-4B
AXIAL-FLOW TURBOJET ENGINE. Carl L. Meyer
and Harry E. Bloomer. November 23, 1948. 60p.
diagrs., photo. (NACA RM E8J25a)

An altitude-wind-tunnel investigation of the perform-
ance and windmilling drag characteristics of an orig-
inal and a modified turbojet engine was conducted
over a range of simulated flight conditions. Per-
formance variables depending upon fuel consumption
that are obtained from data at one altitude cannot be
used to predict these variables at other altitudes;
however, thrust and air-flow values can be predicted
for a limited range of altitudes from data taken at
one altitude. The maximum net thrust of the modi-
fied engine was 3 to 20 percent greater than that of
the original engine. The windmilling engine speed
and drag of the two engines were comparable.


NACA RM E8J25d

INVESTIGATION OF PERFORMANCE OF FURBO-
JET ENGINE WITH CONSTANT- AND VARIABLE-
AREA EXHAUST NOZZLES. Lewis E. Wallner.
November 26, 1948. 43p. diagrs., photos., tab.
(NACA RM E8J25d)


Presents performance of turbojet engine with con-
stant- and variable-area exhaust nozzles at simu-
lated altitudes from 5000 to 45,000 feet and simu-
lated flight Mach numbers from 0.12 to 0.94. The
efficiency of the variable-area nozzle investigated


NACA
RESEARCH ABSTRACTS NO. 57
N.
was 1.5 to 8 percent lower than the efficiency of the--
constant -area nozzle. As a result, lower thrusts
and specific fuel consumption were obtained with
the variable-area nozzle as compared to the
constant-area nozzle.



NACA RM E8J25e

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
TAIL-PIPE BURNING WITH A WESTINGHOUSE
X24C-4B AXIAL-FLOW TURBOJET ENGINE.
William A. Fleming and Lewis E. Wallner.
December 13, 1948. 56p. diagrs., photos., tab.
(NACA RM E8J25e)


Thrust augmentation of a 3000-pound-thrust axial-
flow type turbojet engine by burning fuel In the tail
pipe has been investigated in the NACA Cleveland
altitude wind tunnel. The highest net-thrust increase
obtained in the investigation was 86 percent with a net
thrust specific fuel consumption of 2.91 and a total
fuel-air ratio of 0.0523. The highest combustion
efficiencies obtained with each of the four configura-
tions investigated ranged from 0.71 to 0.96.




NACA RM E8J28

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
WESTINGHOUSE 19B-2, 19B-8, AND 19XB-1 JET-
PROPULSION ENGINES. I OPERATIONAL CHAR-
ACTERISTICS. William A. Fleming. 21p. diagrs.,
photos. (NACA RM E8J28)


Operational characteristics of the 19B-2, 19B-8, and
19XB-1 turbojet engines discussed include effects of
altitude and airspeed on operating range, combustion
stability, starting, acceleration, and functioning of
fuel-control system. Operation of 19B engines above
an altitude of 17,000 feet was limited because of com-
bustion blow-out; operational range of the 19XB-1
engine was slightly better than that of 19B engines.
Starting characteristics of the 19XB-1 engine were
satisfactory but 19B engines did not start
consistently.




NACA RM E8J28a

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
WESTINGHOUSE 19B-2, 19B-8, AND 19XB-1 JET-
PROPULSION ENGINES. II ANALYSIS OF TUR-
BINE PERFORMANCE OF 19B-8 ENGINE. Richard
P. Krebs and Frank L. Suozzi. November 24, 1948.
27p. diagrs., photos. (NACA RM E8J28a)


The performance of the turbine of the 19B-8 turbojet
engine was determined from an investigation of the
complete engine in the altitude wind tunnel over a
range of simulated flight conditions. The effects of
altitude, flight Mach number, and tail-cone position
on turbine performance are discussed. Turbine ef-
ficiency was unaffected by changes in altitude up to
15,000 feet but was a function of tail-cone position
and flight Mach number.







NACA
RESEARCH ABSTRACTS NO. 57

NACA RM E8J28b

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
WESTINGHOUSE 19B-2, 19B-8, AND 19XB-1 JET-
PROPULSION ENGINES. III PERFORMANCE AND
WINDMILLING DRAG CHARACTERISTICS. William
A. Fleming and Robert 0. Dietz, Jr. November 26,
1948. 71p. diagrs., photos. (NACA RM E8J28b)


Performance of the 19B-8 and 19XB-1 turbojet en-
gines and windmilling drag of the 19B-8 engine were
determined in the altitude wind tunnel for a range of
simulated flight conditions. Performance variables
Involving fuel consumption that are obtained from
data at one altitude cannot be used to predict these
variables at altitudes above 15,000 feet; thrust and
air-flow values can be predicted for the range of al-
titudes investigated from data taken at one altitude.
At similar operating conditions, corrected values of
jet thrust and air flow were approximately the same
for both engines.

NACA RM E8J28c

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
WESTINGHOUSE 19B-2, 19B-8, AND 19XB-1 JET-
PROPULSION ENGINES. IV ANALYSIS OF COM-
PRESSOR PERFORMANCE. Robert 0. Dietz and
John K. Kuenzig. November 26, 1948. 48p. diagrs.,
photos. (NACA RM E8J28c)


Performance of the compressors of the 19B-8 and
19XB-1 turbojet engines was determined from an in-
vestigation of the complete engines over a range of
simulated flight conditions. For the range of com-
pressor operation investigated, Reynolds number had
no measurable effect on compressor performance.
The operating lines of the 19B-8 compressor were
on the low-air-flow side of the region of maximum
efficiency. The pressure ratio of the 19XB-1 com-
pressor was higher than that of the 19B-8 compres-
sor.

NACA RM E8J28d

ALTITUDE-WIND-TUNNEL INVESTIGATION OF
WESTINGHOUSE 19B-2, 19B-8, AND 19XB-1 JET-
PROPULSION ENGINES. V COMBUSTION-
CHAMBER PERFORMANCE. Bern rose Boyd.
November 26, 1948. 41p. diagrs., photo. (NACA
RM E8f28d)


Combustion-chamber pressure losses of the 19B-2
and 19B-8 turbojet engines and combustion efficien-
cies of these two engines and the 19XB-1 turbojet
engine are presented. Data were obtained from in-
vestigation of the complete engine over a range of
simulated flight conditions. Maximum combustion
efficiency was shifted to higher engine speeds by in-
creasing altitude. Combustion efficiency was de-
creased at low engine speeds by increasing tunnel
Mach number and was decreased at high engine
speeds by extension of the tail cone.


NACA RM E8J29

SIMULATED ALTITUDE PERFORMANCE OF COM-
BUSTOR OF WESTINGHOUSE 19XB-1 JET-
PROPULSION ENGINE. J. Howard Childs and
Richard J. McCafferty. November 30, 1948. 44p.
diagrs., photo., 2 tabs. (NACA RM E8J29)


17

A 19XB-1 combustor was operated under conditions
simulating zero-ram operation of the 19XB-1 turbo-
jet engine at various altitudes and engine speeds.
The combustion efficiencies and the altitude oper-
ational limits were determined; data were also
obtained on the character of the combustion, the
pressure drop through the combustor, and the com-
bustor outlet temperature and velocity profiles.


NACA RM E8K22

CARBON DEPOSITION OF 19 FUELS IN AN
ANNULAR TURBOJET COMBUSTOR. Jerrold D.
Wear and Edmund R. Jonash. February 3, 1949.
21p. diagrs., 2 tabs. (NACA RMh E8K22)


Effects of fuel properties and change in simulated en-
gine operating conditions on carbon deposition were
evaluated in an annular turbojet combustor with a
diameter of 10-3,8 inches. The fuels included hydro-
carbons of the paraffinic, olefinic, and aromatic
types as well as fuel mixtures. Carbon deposition
increased with increase in boiling temperature of
fuels of the same hydrocarbon type. Aromatic fuels
deposited more carbon than the other types of fuel of
the same boiling temperature. An empirical correla-
tion of the carbon deposition and the boiling tempera-
ture and hydrogen-carbon weight ratio of the fuel was
obtained.


NACA RM E8LO2

COMPARISON OF FLIGHT PERFORMANCE OF
AN-F-58 AND AN-F-32 FUELS IN J35 TURBOJET
ENGINE. Loren W. Acker and Kenneth S.
Kleinknecht. April 7, 1949. 15p. diagrs., photo.,
2 tabs. (NACA RM E8L02)


Performance of fuels corresponding to specifications
AN-F-58 and AN-F-32 was investigated in a J35 tur-
bojet engine in flight. Comparatively, AN-F-58 fuel
indicated slightly higher net thrust and fuel consump-
tion, inferior blowout limits, and similar altitude
starting characteristics and carbon deposits.


NACA RM E8L20a

ALTITUDE PERFORMANCE OF AN-F-58 FUELS
IN J35-C-3 SINGLE COMBUSTOR. Edward G.
Stricker and Warren D. Rayle. June 14, 1949. 21o.
diagrs., 2 tabs. (NACA RM E8L20a)


Three fuel blends conforming to AN-F-58 specifica-
tions were tested in order to determine the influence
of boiling temperatures and aromatic content on alti-
tude performance in single combustor of 4000-pound-
thrust turbojet engine. At simulated engine condi-
tions from an altitude of 30,000 to an altitude of
60,000 feet and 85-percent rated engine speed, the
three AN-F-58 fuels showed little difference in
performance.


NACA RM E9B02

INVESTIGATION OF SEVERAL CLAMSHELL
VARIABLE-AREA EXHAUST NOZZLES FOR TUR-
BOJET ENGINES. Bruce T. Lundin. May 26, 1949.
52p. diagrs., photos. (NACA RM E9B02)











Efficiency of five different types of clamshell varia-
ble area exhaust nozzle for turbojet engines is com-
pared with performance of various conventional
fixed-area conical nozzles. Variable-area nozzles
had satisfactory mechanical reliability even under
afterburning conditions and three of the nozzles in-
vestigated provided jet thrust within 0 to 1-1/2 per-
cent of that obtained with the fixed-area nozzles.
Larger thrust losses for the other two nozzles were
attributed principally to a nonplanar outlet configura-
tion of the movable nozzle flaps.

NACA RM E9C15

TURBINE-ROTOR-BLADE DESIGNS BASED ON
ONE-DIMENSIONAL-FLOW THEORY. I PER-
FORMANCE OF SINGLE-STAGE TURBINE HAVING
40-PERCENT REACTION. Robert E. English and
Cavour H. Hauser. June 10, 1949. 31p. diagrs.,
photo. (NACA RM E9C15)


The performance of the first of a family of turbine-
rotor-blade designs is presented. The rotor blade
was designed for a total-to-static pressure ratio of
4.00 and 40-percent reaction using one-dimensional-
flow theory. Constant static pressure was assumed
over the blade height at the entrance and the exit of
each row of blades. The brake internal efficiency
had a maximum value slightly greater than 0.84 and
was equal to 0.82 at design conditions. For design
conditions, the static pressure varied 5 percent over
the blade height at the rotor exit.

NACA RM E9D01

EVALUATION OF PISTON-TYPE GAS-GENERATOR
ENGINE FOR SUBSONIC TRANSPORT OPERATION.
A. F. Lietzke and Hugh M. Henneberry. July 15,
1949. 29p. diagrs., tab. (NACA RM E9D01)


A piston-type gas-generator engine was evaluated by
comparing performances of a transport airplane
powered by this engine with an airplane having same
characteristics powered by other types of engine.
Turbojet, turbine-propeller, compound, and turbo-
supercharged reciprocating engine were compared
with gas-generator engine. Comparison is based on
pay-load ton-miles per hour operation per ton take-
off gross weight. Gas-generator engine was found to
have a marked advantage over other engine types for
long range at subsonic flight speeds. At short ranges
no one engine showed great superiority, but the turbo-
jet engine exhibits some advantages at high velocity.


NACA RM E9E16

EFFECT OF TEMPERATURE ON PERFORMANCE
OF SEVERAL EJECTOR CONFIGURATIONS. H. D.
Wilsted, S. C. Huddleston and C. W. Ellis. June 13,
1949. 27p. diagrs. (NACA RM E9E16)


An investigation was made to determine the effect of
the primary-jet temperature on the performance of
several ejector configurations. In general, for ejec-
tors with short straight mixing lengths and short
spacings, weight-flow ratio varies directly as the
square root of ratio of total primary temperature to
total secondary temperature. For ejectors with
longer lengths and spacings, which pump greater
weight flows through the secondary annulus, however,


NACA
RESEARCH ABSTRACTS NO. 57

the pressure increase in the plane of the primary---
nozzle exit decreases secondary weight flow so much
that the temperature-ratio correction factor doesn't
completely describe temperature effect on ejector
performance.


NACA RM E9G01

PROPERTIES OF CERTAIN INTERMETALLICS AS
RELATED TO ELEVATED-TEMPERATURE APPLI-
CATIONS. I MOLYBDENUM DISILICIDE. W. A.
Maxwell. October 6, 1949. 27p. diagrs., 4 tabs.
(NACA RM E9G01)


Methods were developed for the preparation, purifi-
cation, and formation by pressing and sintering of
molybdenum ctisilicide. The modulus-of-rupture
strength and oxidation resistance at 20000 and
24000 F were studied as well as the general proper-
ties. Sintered molybdenum disilicide was found to
have high comparative strength at 24000 F and to
possess exceptional resistance to oxidation at high
temperatures. The material deformed plastically
at temperatures well below the melting point despite
complete brittle behavior at room temperature.



NACA RM E9G08

ALTITUDE PERFORMANCE AND OPERATIONAL
CHARACTERISTICS OF 29-INCH-DIAMETER TAIL-
PIPE BURNER WITH SEVERAL FUEL SYSTEMS
AND FLAME HOLDERS ON J35 TURBOJET ENGINE.
E. William Conrad and William R. Prince.
November 8, 1949. 50p. diagre., photos., tab.
(NACA RM E9G08)


An investigation of afterburning was conducted in the
Lewis altitude wind tunnel using a full-scale turbojet
engine to obtain information on afterburner design
variables. Radial distribution and direction of injec-
tion of the tail-pipe fuel and several flame-holder
types were studied. Direction of fuel injection had no
effect on combustion efficiency; however, poor radial
fuel distribution resulted in low combustion efficien-
cies. As altitude was increased, the decrease in
peak combustion efficiency became more rapid as the
blocking area of the flame holder was reduced.
Problems of tail-pipe cooling and ignition are dis-
cussed and satisfactory solutions obtained.




NACA RM E9G13

CYCLIC ENGINE OPERATION OF CAST VITALLIUM
TURBINE BLADES AT AN EXHAUST-CONE GAS
TEMPERATURE OF 1440 200 F. Charles Yaker
and Floyd B. Garrett. September 19, 1949. 41p.
diagrs., photos., 4 tabs. (NACA RM E9G13)


An investigation was conducted to study engine per-
formance of cast Vitalliumn turbine blades at an es-
timated maximum blade temperature of 16000 F.
The blades, mounted in a Timken 16-25-6 wheel,
were subjected to 20-minute cycles consisting of
approximately 5 minutes at idle and 15 minutes at
rated speed. After 7-1 '2 cycles (2 hr, 30 min), one








NACA
RESEARCH ABSTRACTS NO. 57

blade fractured due to fatigue, a large number of
blades had Intercrystalline cracks, and two blades
had large transcrystalline cracks. Examination of
broken and unbroken blades disclosed no correlation
between blade failure and elongation as compared
with grain size and hardness.

NACA RM E9G25

VIBRATIONAL MODES OF SEVERAL HOLLOW TUR-
BINE BLADES AND OF SOLID TURBINE BLADE OF
SIMILAR AERODYNAMIC DESIGN. R. H. Kemp and
J. Shifman. October 3, 1949. 17p. diagrs., photos.
(NACA RM E9G25)


Vibrational modes of several hollow turbine blades
and a solid turbine blade of similar aerodynamic de-
sign were experimentally determined. Most of the
observed vibrational modes of the hollow blades could
be excited if the blades were operated in a conven-
tional turbojet engine. A breathing effect was found
in hollow blades at approximately 1230 cycles per
second in which the blade walls moved toward each
other and then away, producing stress concentrations
at the leading and trailing edges. Little similarity in
vibrational modes existed between the hollow blades
and the solid blade. Approximately twice as many
readily excited modes were found in the hollow
blades as in the solid blade.


NACA RM E9107

VIBRATION OF LOOSELY MOUNTED TURBINE
BLADES DURING SERVICE OPERATION OF A
TURBOJET ENGINE WITH CENTRIFUGAL COM-
PRESSOR AND STRAIGHT-FLOW COMBUSTION
CHAMBERS. W. C. Morgan, R. H. Kemp and S. S.
Manson. November 3, 1949. 18p. diagrs., photos.
(NACA RM E9I07)


Vibration characteristics of loosely mounted turbine
blades were determined during service operation of a
turbojet engine. High-temperature strain gages were
used to measure turbine-blade vibrations. Vibration
occurred in first bending and first torsional modes;
in addition, a small number of complex-mode vibra-
tions was observed. Comparison was made between
vibrations in loosely mounted blades and those ob-
served during a previous Investigation of similar
blades tightly mounted in a turbine wheel. The com-
parison did not indicate that any considerable gain in
damping was effected by the use of loosely mounted
blades.


NACA RM E9I23

ALTITUDE-CHAMBER PExFORMANCE OF
BRITISH ROLLS-ROYCE NENE II ENGINE. I -
STANDARD 18.75-INCH-DIAMETER JET NOZZLE.
Zelmar Barson and H. D. Wilsted. September 23,
1949. 58p. diagrs., 2 tabs. (NACA RM E9123)

An altitude-chamber investigation of the British
Rolls-Royce Nene II turbojet engine was conducted
over a range of altitudes from sea level to 60,000
feet and ram-pressure ratios from 1.00 to 3.50.
Decreasing compressor pressure ratio and efficiency
at high altitude prevented prediction of engine per-
formance parameters for altitudes above 20,000


19

feet from data obtained at any one altitude. Engine
performance generalized to a single curve for ram-
pressure ratio for all conditions when critical flow
existed in the jet nozzle.


NACA RM E9K04

CORRELATION OF LABORATORY SMOKE TEST
WITH CARBON DEPOSITION IN TURBOJET COM-
BUSTORS. Arthur M. Busch. February 3, 1950.
15p. diagrs., 3 tabs. (NACA RM E9K04)


A function of flame heights at the sooting point in a
simple wick lamp and boiling point of 19 fuels was
correlated with carbon deposition in a 10-3/8 inch
annular turbojet combustor operating at a single set
of conditions. Two similar Investigations with
different combustors, operating conditions, and mne
different fuels yielded similar correlations. The
simple wick lamp may be of value In the evaluation
of turbojet fuels.


NACA RM E9L05

EFFECTS OF OBSTRUCTIONS IN COMPRESSOR
INLET ON BLADE VIBRATION IN 10-STAGE AXIAL-
FLOW COMPRESSOR. Andre J. Meyer, Jr., Howard
F. Calvert and C. Robert Morse. February 13,
1950. 16p. diagrs. (NACA RM E9L05)

Resistance-wire strain gages were used to measure
blade-vibration characteristics in 10-stage axial-
flow compressor during normal jet-engine operation.
Effects produced by obstructions in path of air flow
entering compressor were also investigated. Aero-
dynamic damping of vibrating blades was evaluated.
Results indicated that obstructions affected blade vi-
brations throughout compressor, that obstructions
could be so located as to reduce some vibrations, and
that aerodynamic damping accounted for about four-
fifths of total dynamic blade damping.



NACA RM L6JO9

HIGH-SPEED WIND-TUNNEL TESTS OF A 1, 16-
SCALE MODEL OF THE D-558 RESEARCH AIR-
PLANE. LIFT AND DRAG CHARACTERISTICS OF
THE D-558-I AND VARIOUS WING AND TAIL CON-
FIGURATIONS. John B. Wright and Donald L.
Loving. April 18, 1947. 43p. diagrs., 2 tabs.
(NACA RM L6J09)


Test results indicated the airplane will have satis-
factory lift and drag characteristics up to and in-
cluding the design Mach number of 0.85. The swept-
wing and low-aspect-ratio wing configurations showed
pronounced improvement in maintaining lift through-
out the Mach number range tested and in increasing
critical Mach numbers to the order of 0.9.


NACA RM L6K08

RESUME OF WIND-TUNNEL DATA ON THE EF-
FECT OF EXTERNAL STORES ON STABILITY OF
MODELS OF MILITARY AIRPLANES. H. Norman
Silvers and Raymond D. Vogler. December 19,
1946. 7p. 2 tabs. (NACA RM L6K08)









20

A summary, in tabular form, is presented on the ef-
fect on static stability of tanks, torpedoes. bombs,
and radar domes. The data obtained indicate that at
Mach numbers below 0.4 the effects of external
stores on static longitudinal stability were small but
may not be negligible. Although data were meager
for Mach numbers above 0.4. results indicated that
wing-tip tanks which were well faired to the wing
contour caused reasonably small changes in static
longitudinal stability. The available data were in-
sufficient to estimate the general effects of external
stores on lateral stability.



NACA RM L6K08b

PRELIMINARY TESTS OF A BURNER FOR RAM-
JET APPLICATIONS. Paul W. Huber. January 15,
1947. 14p. diagrs. (NACA RM L6K08b)


Preliminary tests of a small ram-jet burner have
been made and tests of this burner indicate efficient
combustion characteristics (combustion was 80 per-
cent complete) and low aerodynamic losses at the
conditions tested. The fact that the measured com-
bustion was so nearly complete is due mainly to the
method of fuel injection and to mixing of fuel and air
before ignition. The pressure drop due to frictional
losses at the burner at the inner passage of the ram
jet is small, since the air is not greatly disturbed to
obtain mixing and ignition.



NACA RM L6L19

THE EFFECT OF HIGH SOLIDITY ON PROPELLER
CHARACTERISTICS AT HIGH FORWARD SPEEDS
FROM WIND-TUNNEL TESTS OF THE NACA
4-(3)(06.3)-06 AND NACA 4-t3)t06.4)-09 TWO-
BLADE PROPELLERS. James B. Delano.
February 27, 1947. 50p. diagrs., photos. (NACA
RM L6L19)


Two two-blade propellers were tested in the Langley
8-foot high-speed tunnel through a range of blade
angle from 200 to 700 for free-stream Mach numbers
from 0.165 to 0.725 to determine the effects of high
solidity and compressibility on propeller character-
istics.



NACA RM L6L27

EFFECTS OF A FUSELAGE AND VARIOUS HIGH-
LIFT AND STALL-CONTROL FLAPS ON AERO-
DYNAMIC CHARACTERISTICS IN PITCH OF AN
NACA 64-SERIES 400 SWEPT-BACK WING
D. William Conner and Robert H. Neely. May 26,.
1947. 40p. diagrs., photos., tab. (NACA
RM L6L27)


Wind-tunnel tests were made to determine low-speed
lift, drag, and pitching-moment characteristics.
The wing had an aspect ratio of 4 and a taper ratio
of 0.625. Low-, middle-, and high-wing-fuselage
combinations were tested at Reynolds numbers of
3 x 106 and 8.1 x 106.


NACA
RESEARCH ABSTRACTS NO. 57
'S
NACA RM L7C11 '

AERODYNAMIC CHARACTERISTICS OF A 450
SWEPT-BACK WING WITH ASPECT RATIO OF 3.5
AND NACA 25-50105)-50(05) AIRFOIL SECTIONS.
Anthony J. Proterra. August 4. 1947. 21p.
diagrs., photo. (NACA RM L7C11)


Scale effects were investigated at Reynolds numbers
ranging from 2.1 x 106 to 8.0 x 106; the effects of
yaw were investigated at a Reynolds number of
4.1 x 106. Results indicate that the wing has poor
characteristics from low-speed considerations.


NACA RM L7C24a

INVESTIGATION OF INTAKE DUCTS FOR A HIGH-
SPEED SUBSONIC JET-PROPELLED AIRPLANE.
Herbert N. Cohen. April 23, 1947. 24p. diagrs.,
photos., tab. (NACA RM L7C24a)


Results of pressure-loss measurements are pre-
sented for full-scale models of two ducts under con-
sideration for use in the induction system of an ex-
perimental jet-propelled airplane. Supplementary
pressure-loss measurements were made on the
better of the two ducts, designated duct II, first with
carborundum grains in the duct inlet and then with a
spoiler in order to obtain an indication of the impor-
tance of inlel roughness and surface discontinuities.
Additional measurements were made of duct II in-
corporating a horizontal "splitter" vane which was
under consideration for structural reasons.


NACA RM L7D03

AN INVESTIGATION OF THE LOW-SPEED CHAR-
ACTERISTICS OF TWO SHARP-EDGE SUPERSONIC
INLETS DESIGNED FOR ESSENTIALLY EXTERNAL
SUPERSONIC COMPRESSION. John S. Dennard.
June 6, 1947. 32p. diagrs.. photos. (NACA
RM L7D03)

The investigation has been conducted at lowairspeeds
to obtain preliminary information concerning the
surface-pressure, drag, and pressure recovery of
such inlets. Surface-pressure measurements and
surveys of the pressures in the internal andexternal
flow were obtained at angles of attack of 00 and 60
for a wide range of inlet-velocity ratio.



NACA RM LTDI4

AERODYNAMIC CHARACTERISTICS OF A 420
SWEPT-BACK WING WITH ASPECT RATIO 4 AND
NACA 641-112 AIRFOIL SECTIONS AT REYNOLDS
NUMBERS FROM 1,700,000 TO 9,500,000.
Robert H. Neely and D. William Conner. May 23,
1947. 39p. diagrs., photos. (NACA RM L7D14)


Wind-tunnel tests were made to determine low-speed
aerodynamic characteristics in pitch and in yaw at
high Reynolds numbers. The characteristics In
pitch were obtained over a Reynolds number range
from 1.7 x 106 to 9.5 x 106 and the characteristics
in yaw from 1.7 x 106 to 5.3 x 106.








NACA
RESEARCH ABSTRACTS NO. 57


NACA RM L7E12

FURTHER INVESTIGATION OF NACA 4-(5)(08)-03
TWO-BLADE PROPELLER AT HIGH FORWARD
SPEEDS. Melvin M. Carmel and Harold L.
Robinson. May 26, 1947. 46p. diagrs. (NACA
RM L7E12)


Tests of an NACA two-blade propeller have been
made in the Langley 8-foot high-speed tunnel for
blade angles of 450 and 600 extending the Mach num-
ber range from that of previous tests of this pro-
peller up to a forward Mach number of 0.913.



NACA RM L7E13

EFFECTS OF A FUSELAGE ON THE AERODYNAM-
IC CHARACTERISTICS OF A 420 SWEPTBACK
WING AT REYNOLDS NUMBERS TO 8.000,000.
Reino J. Salmi, D. William Conner and Robert R.
Graham. June 10, 1947. 32p. diagrs., photos
(NACA RM L7E13'


Tests were made in pitch at Reynolds numbers oi
3.040.000 and 8,090,000 and in yaw at Reynolds nun;-
bers of 1,720,000 and 5.300,000. The wing had an
aspect ratio of 4. a taper ratio of 0.625 and NACA
641-112 airfoil sections. Tests were made with the
wing in high-, los -, and midwing positions Tests
were made without flaps and with 18.4-percent -chord
split flaps extending from 12.3 to 50 percent of the
semispan. The presence ol the fuselage had little
effect on the values of the maximum lift coefficient
and the slope of the lift curve, but caused a destabi-
lizing shift in the rate of change of pitching moment
with lift.



NACA RM L7E23

LOW-SPEED CHARACTERISTICS IN PItCH OF A
420 SWEPTBACK WING WITH ASPECT RATIO 3.9
AND CIRCULAR-ARC AIRFOIL SECTIONS.
Robert H. Neely and William Koven. November 13.
1947. 42p. diagrs., photos., 2 tabs. (NACA
RM L7E23)


Characteristics of basic wing were poor. The ef-
fects on the wing characteristics of extensible
round-nose leading-edge flaps located on the out-
board 70 percent of the semispan, and of a fuselage
with fineness ratio 10.2 located in the low, middle,
and high positions were investigated. I its covered
a range of Reynolds numbers from 3.09 z 106 to
9.60 x 106.



NACA RM L7E29

ESTIMATION OF RANGE OF STABILITY DERIVA-
TIVES FOR CURRENT AND FUTURE PILOTLESS
AIRCRAFT. Marvin Pitkin and Herman 0.
Ankenbruck. October 8, 1947. 22p. diagrs., tab.
(NACA RM L7E29)


21


The estimated range of the aerodynamic and mass
stability derivatives for airplanes and missiles Is
given primarily as an aid to designers of "flight
tables" and "simulators."

NACA RM L7FO4a

LODW-SPEED CHARACTERISTICS IN PITCH OF A
340 SWEPTFORWARD WING WITH CIRCULAR-ARC
AIRFOIL SECTIONS. D. William Conner and
Patrick A. Cancro. January 9, 1948, 35p. diagrs.,
photos. (NACA RM L7FO4a)


Contains low-speed lift, drag, and pitching-moment
characteristics of a 340 sweptforward wing with as-
pect ratio 3.9 and circular-arc airfoil sections tested
with a fuselage and high-lilt and stall-control de-
vices. Adding a fuselage to the basic wing reduced
the undesirable variations in pitching moment in the
high-lift range. Adding full-span, extended, round-
nose, leading-edge flaps to the midwing-fuselage
combination increased the maximum lift coefficient
to 1.30 without seriously decreasing the longitudinal
stability. Combinations with half-span split flaps
had unfavorable pitching-moment characteristics.


NACA RM L7F13

MEASUREMENTS OF AERODYNAMIC CHARACTER-
ISTICS OF A 350 SWEPTBACK NACA 65-009 AIR-
FOIL MODEL WITH 1. 4-CHORD PLAIN FLAP BY
THE NACA WING-FLOW METHOD. Harold I.
Johnson. August 5, 1947. 72p. diagrs., photos.
(NACA RM L7F13)


As part of a general investigation of the stability and
control characteristics of various airfoil-flap com-
binations in the transonic speed range, measurements
were made by the NACA wing-flow method of the
lift, pitching-moment, and hinge-moment character-
istics of a 350 sweptback NACA 65-009 airfoil of
aspect ratio 3.04, with a full-span 1. 4-chordunsealed
plain flap. The tests covered Mach numbers from
0.55 to 1.10, Reynolds numbers from about 500,000
to 1,300,000, angles of attack from -2o to 150, and
fap deflections from about -200 to 200.


NACA RM L7F16

LANGLEY FULL-SCALE-TUNNEL INVESTIGATION
OF MAXIMUM LIFT AND STABILITY CHARACTER-
ISTICS OF AN AIRPLANE HAVING APPROXIMATE -
LY TRIANGULAR PLAN FORM (DM-1 GLIDER).
J. Calvin Lovell and Herbert A. Wilson, Jr.
August 5, 1947. 44p. diagrs., photos., tab.
(NACA RM L7F16)


This glider had an aspect ratio of 1.8 and 600 sweep-
back. The investigation consisted of the determina-
tion of the separate effects of the following modifi-
cations made to the glider on its maximum lift and
stability characteristics: (a) installation of sharp
leading edges over the inboard semispan of the wing,
(b) removal of the vertical fin. (c) sealing of the
elevon control-balance slots, (d) installation of re-
designed thin vertical surfaces, (e) installation of
faired sharp leading edges, and (f) installation of
canopy.








22

NACA RM L7F19

PRELIMINARY INVESTIGATION OF SPOILER
LATERAL CONTROL ON A 420 SWEPTE.ACK WING
AT TRANSONIC SPEEDS. Leshlie E. Schneiter and
Howard L. Ziff. August 12, 1947. 13p. diagrs.
(NACA RM L7F19)


This investigation was performed on a double-wedge
type semispan wing by the NACA wing-flow method
of testing. Above a Mach number of 0.6, the effec-
tiveness of the spoiler in producing rolling moment
at an angle of attack of 30 increased as a Mach num-
ber of 0.90 was approached. At Mach numbers
greater than 0.90, the spoiler effectiveness de-
creased rapidly until a Mach number of 1.05 was at-
tained, above which point the effectiveness increased
slightly as speed was further increased.

NACA RM L7124

AN ANALYSIS OF DUCTED-AIRFOIL RAM JETS
FOR SUPERSONIC AIRCRAFT. Paul R. Hill and
A. A. GammaL July 7, 1948. 43p. diagrs. (NACA
RM L7124)

The effect of sweepback on the external aerodynamics
of ducted airfoils with the leading edge ahead of the
Mach cone is shown to be small. Possible total pres-
sure recoveries are determined for two-dimensional
wedge-type diffuser inlets in order to compute thrust
and propulsive coefficients for a wide range of hydro-
carbon fuel-air ratios. The possible range and ac-
celeration performances are determined for aircraft
with fineness ratio 10, parabolic body-of-revolution
fuselages with both ducted wing and ducted tails of
various sizes relative to the fuselage size.


NACA RM L7130

YAW CHARACTERISTICS AND SIDEWASH ANGLES
OF A 420 SWEPTBACK CIRCULAR-ARC WING WITH
A FUSELAGE AND WITH LEADING-EDGE AND
SPLIT FLAPS AT A REYNOLDS NUMBER OF
5,300,000. Reino J. Salmi and James E. Fitzpatrick.
December 10, 1947. 38p. diagrs., photos., tab.
(NACA RM L7130)


Contains results of tests of a 420 sweptback wing at
a Mach number of 0.11. The dihedral effect of the
sharp-edged-airfoil wing became negative at moder-
ate lift coefficients similar to low Reynolds number
results of round-nosed airfoil wings. Flap deflec-
tion produced positive effective dihedral. The plain
wing had neutral directional stability with flaps neu-
tral but was stable with flaps deflected. Air-stream
surveys of the sidewash angles of various wing fuse-
lage combinations showed that maximum vertical-
tall effectiveness would be obtained on a low-wing
combination.



NACA RM L7J03

INVESTIGATION OF PRESSURE DISTRIBUTION
OVER AN EXTENDED LEADING-EDGE FLAP ON A
420 SWEPTBACK WING. D. William Conner and
Gerald V. Foster. December 19, 1947. 12p. diagrs.
(NACA RM L7J703)


NACA
RESEARCH ABSTRACTS NO. 57

Pressure distributions over an extended leadlg-
edge flap, mounted on a 42o sweptback wing and -
deflected 500, are presented for several angles of
attack. Results indicate that the flap normal-force
coefficient increased almost linearly with angle of
attack to a maximum value of 3.25. Peak negative
pressures built up at the flap leading edge with angle
of attack and caused the chordwise location of the
flap center of pressure to be moved forward. For
the high angle-of-attack range, the center of pres-
sure ranged between 49 and 55 percent of the flap
chord.

NACA RM L7J10

PRELIMINARY TESTS TO DETERMINE THE MAX-
IMUM LIFT OF WINGS AT SUPERSONIC SPEEDS.
James J. Gallagher and James N. Mueller.
December 11, 1947. 41p. diagrs., photos., 3 tabs.
(NACA RM L7J10)


Contains a variety of wing plan forms of random
thickness distribution which were tested at Mach
numbers of 1.55, 1.90, and 2.32 at angles of attack
ranging from zero up through the angle at which max-
imum lift occurred. In general, at these Mach num-
bers the value of maximum lift coefficient was ap-
proximately 1.05 0.05; it appeared to be Independent
of plan form and decreased slightly with increasing
Mach number. No discontinuities In lift occurred
from zero angle of attack through maximum lift,
which was attained at approximately 400 angle of at-
tack. Lift-drag ratios at maximum lift were of the
order of 1.0.

NACA RM L7K13

WIND-TUNNEL INVESTIGATION OF THE LOW-
SPEED STABILITY AND CONTROL CHARACTER-
ISTICS OF A MODEL WITH A SWEPTBACK VEE
TAIL AND A SWEPTBACK WING. Edward C.
Polhamus. May 25, 1948. 31p. diagrs. (NACA
RM L7K13)

Tests were made in the Langley 300-mph 7- by 10-
foot tunnel of a complete model with a sweptback vee
tail and a sweptback wing to determine its low-speed
stability and control characteristics. Tests were
also made of the wing alone and of the wing in com-
bination with the fuselage. Comparisons are made
with the results of tests of the same tall panel with
zero dihedral (horizontal tail) on the same wing-
fuselage combination.


NACA RM L8B13

MEASUREMENT THROUGH THE SPEED OF SOUND
OF STATIC PRESSURES ON THE REAR OF UN-
SWEPT AND SWEPTBACK CIRCULAR CYLINDERS
AND ON THE REAR AND SIDES OF A WEDGE BY
THE NACA WING-FLOW METHOD. Richard H.
Sawyer and Fred L. Daum. July 21, 1948. 13p.
diagrs., photos. (NACA RM L8B13)


Static-pressure measurements were made through
the speed of sound by the NACA wing-flow method at
the rear of two unswept circular cylinders of differ-
ent length-diameter ratio and one 450 sweptback
circular cylinder. Additional measurements were
made at the rear and sides of a wedge. A Mach num-
ber range from about 0.7 to about 1.2 was covered
in the tests.








NACA
RESEARCH ABSTRACTS NO. 57

NACA RM L8D23

TESTS OF THE NACA 641-012 AND 641A012 AiR-
FOILS AT HIGH SUBSONIC MACH NUMBERS.
W. F. Uindey and Milton D. Humphreys. July 9,
1948. 19p. diagram tab. (NACA RM L8D23)


An investigation of NACA 64i-012 and 641A012 air-
foils at Mach numbers between 0.35 and 0.89 and at
low Reynolds numbers showed the effect of increas-
ing the trailing-edge angle from 9o (NACA 641-012)
to 140 (NACA 641A012).



NACA RM L8H25

STATIC LONGITUDINAL AERODYNAMIC CHARAC-
TERISTICS OF A 520 SWEPTBACK WING OF AS-
PECT RATIO 2.88 AT REYNOLDS NUMBERS FROM
2,000,000 TO 11,000,000. James E. Fitzpatrick and
Gerald V. Foster. November 16, 1948. 21p.
diagrs., photo. (NACA RM L8H25)


Presents results of test of a 520 sweptback wing
with aspect ratio of 2.88, taper ratio 0.625, and
NACA 641-112 airfoil sections perpendicular to the
0.282-chord line for Reynolds number range of
2,000,000 to 11,000,000. Tests were made with and
without half-span split flaps for both the smooth and
rough conditions. The tests also included a study of
the flow changes at moderate to high lift coefficients.




NACA RM L8H25a

STATIC TESTS OF FOUR TWO-BLADE NACA PRO-
PELLERS DIFFERING IN CAMBER AND SOLIDITY.
Robert J. Platt, Jr. December 2, 1948. 23p.
diagrs., photo. (NACA RM L8H25a)


Contains static-test data on two-blade propellers of
the NACA 4-(3)(08)-03, 4-(5)(08)-03, 4-(10)(08)-03,
and 4-(3)(08)-045 designs at blade angles from 00 to
400. Some earlier data are included which were ob-
tained with the propellers flattering.




NACA RM L8H30

INVESTIGATION OF THE EFFECT OF SWEEP ON
THE FLUTTER OF CANTILEVER WINGS. J. G.
Barmby, H. J. Cunningham and 1. E. Garrick.
November 15, 1948. 70p. diagrs., photo., 7 Labs.
(NACA RM L8H30. Now issued as TN 2121;
Rept. 1014)


An experimental and analytical investigation of the
flutter of uniform, cantilever, sweptback wings is
reported. The experiments employed groups of
wings swept back by rotating and by shearing. The
angle of sweep ranged from 00 to 600 and the Mach
numbers extended to approximately 0.9. Compar-
ison with experiment indicates that the analysis de-
veloped in the present paper is satisfactory for
nearly uniform cantilever wings of moderate length-
to-chord ratios.


23

NACA RM L8117

A METHOD FOR CALCULATING FLOW FIELDS OF
COWLINGS WITH KNOWN SURFACE-PRESSURE
DISTRIBUTIONS. Robert W. Boswinkle, Jr.
November 22, 1948. 23p. diagrs. (NACA RM L8I17)


Describes a way in which the data of three previous
papers may be utilized to compute the incompressible
flow fields of cowling-spinner combinations and open-
nose inlets at zero incidence. The method consists of
regarding the cowling surface as replaced by a ring
vortex sheet whose strength at any point is equal to
the local tangential velocity. The field of the ring
vortex sheet is Integrated to give the induced veloci-
ties of the body. An approximate method is given for
cowlings at angles of attack. A comparison of calcu-
lated points with experimental data indicates that the
method gives adequate accuracy for propeller design
purposes. The application of the Prandtl-Glauert
rule for compressible flow is given in detail.



NACA RM L8I24

SURFACE-PRESSURE DISTRIBUTIONS ON A
SYSTEMATIC GROUP OF NACA 1-SERIES COWL-
INGS WITH AND WITHOUT SPINNERS. Robert W.
Boswinkle, Jr. and Arvid L. Keith, Jr. November 30,
1948. 188p. diagrs., 3 tabs. (NACA RM L8124)


Presents the static-pressure distributions on the tops
of 79 NACA I -series cowling-spinner combinations
and9NACA 1-series open-nose cowlings over wide
ranges of inlet-velocity ratio at angles of attack of
00, 20, 40, and 60 for use in calculating the flow
fields of such cowlings. This flow field information
is useful in the design of propellers. Some effects of
changes in the internal lip shape and of propeller
operation on the surface pressures on the cowlings
are shown.



NACA RM L8J01

INVESTIGATION OF HORN BALANCES ON A 450
SWEPTBACK HORIZONTAL TAIL SURFACE AT
HIGH SUBSONIC SPEEDS. Harold S. Johnson and
Robert F. Thompson. December 3, 1948. 63p
diagrs., photo., 2 tabs. (NACA RM L8J01)


A wind-tunnel investigation of horn balances on a
450 sweptback, semispan, horizontal tail surface
was made to determine the effects of horn size and
inboard-edge fairing at a Mach number of 0.30 and to
determine the effects of compressibility up to a Mach
number of 0.89. Presented are lift, drag, pitching-
moment, and hinge-moment data and lift and hinge-
moment parameters.



NACA RM L8K19

WIND-TUNNEL INVESTIGATION AT LOW SPEEDS
OF VARIOUS PLUG-AILERON AND LIFT-FLAP
CONFIGURATIONS ON A 420 SWEPTBACK SEMI-
SPAN WING. Leslie E. Schneiter and James M.
Watson. January 26. 1949. 45p. diagrs., photos.
(NACA RM L8K19)








24

Contains results and discussion of a high-lift and
lateral-control investigation at low speed on a 420
sweptback semispan-wing model. The lift charac-
teristics of a full-span slotted flap at various deflec-
tions and positions of the flap nose, a partial-span
slotted flap, and a partial-span Zap flap were deter-
mined in addition to a determination of the lateral-
control characteristics of a plug-aileron configura-
tion and various modifications thereof, and the char-
acteristics of a partial-span plain aileron in both the
flap-retracted and flap-deflected configurations.



NACA RM L8106

AN INVESTIGATION OF THE CHARACTERISTICS
OF THREE NACA 1-SERIES NOSE INLETS AT SUB-
CRITICAL AND SUPERCRITICAL MACH NUMBERS.
Robert E. Pendley and Norman F. Smith. January 13,
1949. 38p. diagrs., photo. (NACA RM L8L06)


Results of measurements of pressure distribution
and drag are presented for NACA 1-65-050, NACA
1-50-100, and NACA 1-40-200 nose inlets. Data
were obtained at two fineness ratios and at Mach
numbers up to 0.925 for a range of Inlet-velocity
ratio. An appreciable margin was measured between
the critical Mach number and the Mach number at
which a drag rise occurred. A sharp pressure peak
raised at the inlet lip at low inlet-velocity ratios did
not affect the Mach number for drag rise.



NACA RM L9A07

LOW-SPEED INVESTIGATION OF AILERON AND
SPOILER CHARACTERISTICS OF A WING HAVING
420 SWEEPBACK OF THE LEADING EDGE AND
CIRCULAR-ARC AIRFOIL SECTIONS AT REYNOLDS
NUMBERS OF APPROXIMATELY 6.0 x 106.
Stanley H. Spooner and Robert L. Woods. March 10,
1949. 58p. diagram photos. (NACA RM L9A07)


The lateral characteristics of an aileron and several
spoiler arrangements on a 420 sweptback wing with
symmetrical circular-arc airfoil sections and vari-
ous high-lift and stall-control devices are shown.
Includes aileron normal-force, hinge-moment, and
balance-chamber pressure coefficients. The
Reynolds number of the tests, which were conducted
in the Langley 19-foot pressure tunnel, varied from
5.3 x 106 to 6.9 x 106 which corresponded to a Mach
number range of 0.11 to 0.15.


NACA RM L9A1B

PRESSURE-DISTRIBUTION MEASUREMENTS
OVER AN EXTENSIBLE LEADING-EDGE FLAP ON
TWO WINGS HAVING LEADING-EDGE SWEEP OF
420 AND 520. Reino J. Salmi. March 7, 1949.
36p. diagrs., photos. (NACA RM L9A18)


Contains results of pressure-distribution measure-
ments over a leading-edge flap on two wings having
leading-edge sweep of 420 and 520 with NACA
641-112 sections. The 420 sweptback wing was of
aspect ratio 4.01, taper ratio 0.625, equipped with
half-span split flaps and was tested in combination
with a circular fuselage at a Reynolds numuei of


NACA
RESEARCH ABSTRACTS NO. 57

5.12 x 106 and Mach number of 0.11. The 520'N
sweptback wing was of aspect ratio 2.88, taper ratidN.
0.625, equipped with half-span split flaps and a
fence. The tests were made at a Reynolds number
of 4.4 x 106 and Mach number of 0.08, and at various
angles of yaw.



NACA RM L9FO10

PRELIMINARY AERODYNAMIC INVESTIGATION OF
THE EFFECT OF CAMBER ON A 600 DELTA WING
WITH ROUND AND BEVELED LEADING EDGES.
John M. Riebe and Joseph E. Fikes. August 16,
1949. 46p. diagrs., photos., tab. (NACA
RM L9F10)


Contains the results of an investigation in the
Langley 300-mph 7- by 10-foot wind tunnel to deter-
mine the effect of camber on lift-drag ratio and the
longitudinal and lateral stability characteristics on
a 600 delta wing at Reynolds numbers of 1.5 x 106
and 3.0 x 106. Angles of attack ranged from 00 to
400; angles of yaw ranged from -40 to 200. Camber
was produced by full-span leading-edge flaps of
round and beveled shapes deflected through a range
of 00 to 600.




NACA RM L9H09

PERFORMANCE CHARACTERISTICS OF TWO 60
AND TWO 120 DIFFUSERS AT HIGH FLOW RATES.
William J. Nelson and Eileen G. Popp. October 19,
1949. 26p. diagrs. (NACA RM L9H09)


The aerodynamic characteristics of two circular-
inlet and two annular-inlet diffusers of fixed area
ratio (1.75) have been determined at inlet Mach num-
bers from 0.2 to the choked condition. Thepressure-
loss coefficient and diffuser effectiveness of each of
these diffusers is shown to be essentially Independ-
ent of Reynolds number in the subcritical-flow range;
the performance falls off rapidly when sonic velocity
is exceeded at any point in the system. Pressure
distributions across both the inlet and exit and along
the diffuser walls are presented. Inlet-boundary-
layer profiles were also measured.




NACA RM L9H23

AN INVESTIGATION OF THE CHARACTERISTICS
OF AN UNSWEPT WING OF ASPECT RATIO 4.01
IN THE LANGLEY 8-FOOT HIGH-SPEED TUNNEL.
Ralph P. Bielat and Maurice S. Cahn. November 8,
1949. 32p. diagrs., 2 tabs. (NACA RM L9H23)


An investigation of the characteristics of a wing of
aspect ratio 4.01 was conducted at high subsonic
Mach numbers in the Langley 8-foot high-speed tun-
nel. The results of this investigation indicate that
the severe changes in aerodynamic characteristics
usually associated with wings of average or high as-
pect ratio were alleviated and were delayed to higher
Mach numbers.







NACA
RESEARCH ABSTRACTS NO.57

NACA RM L9JS03

SURVEY OF TWO-DIMENSIONAL DATA ON
PITCHING-MOMENT CHANGES NEAR MAXIMUM
LIFT CAUSED BY DEFLECTION OF HIGH-LIFT
DEVICES. Jerold M. Bidwell and Jones F. Cahill.
December 2, 1949. 13p. diagram tab. (NACA
RM L9JOS)


The large pitching-moment increments associated
with deflection of certain types of trailing-edge
high-lift devices have made it difficult or impossible
to obtain trim during landing and take-off. As an
aid in the selection of high-lift devices, therefore, a
survey has been made of two-dimensional data on
the changes in pitching moment resulting from de-
flection at various types of high-lift devices. The
types of high-lift devices investigated include ex-
tensible and nonextensible leading-edge and trailing-
edge flaps and slats.




NACA RM E6J21

EXPERIMENTAL INVESTIGATION OF THRUST
AUGMENTATION OF A TURBOJET ENGINE AT
ZERO RAM BY MEANS OF TAIL-PIPE BURNING.
Bruce T. Lundin, Harry W. Dowman and David S.
Gabriel. January 6, 1947. 34p. diagrs., photos.,
tab. (NACA RM E6J21)


The performance of a turbojet engine equipped with a
tailpipe burner designed by the NACA has been inves-
tigated at zero ram over a range of rotor speeds and
tailpipe-burner fuel flows. The burner consisted
essentially of an enlarged tailpipe incorporating fuel-
spray nozzles and a flame holder. An adjustable-
area exhaust nozzle is installed at the burner dis-
charge.




THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL TO
UNCLASSIFIED, 1/8/54.



NACA RM A8A20

AN INVESTIGATION OF SUBMERGED AIR INLETS
ON A 1/4-SCALE MODEL OF A TYPICAL
FIGHTER-TYPE AIRPLANE. Noel K. Delany.
June 2, 1948. 47p. diagrs., photos. (NACA
RM A8A20)

Results are presented for ramp plan forms with
parallel and with diverging walls and show the effect
of the duct-entrance location (forward of the wing and
over the wing), internal ducting efficiency, and de-
flectors. The air inlets having the ramps with di-
verging walls were satisfactory in both locations
tested on the fuselage. The submerged air inlets
with parallel ramp walls had lower ram pressure
recoveries for the normal operating range. The
ram pressure-recovery ratios measured at the in-
lets were higher for the forward location of the in-
lets than for the aft location.


25

NACA RM E7H27

INDIRECT METHODS FOR OBTAINING RAM-JET
EXHAUST-GAS TEMPERATURE APPLIED TO
FUEL-METERING CONTROL. Eugene Perchonok,
William H. Sterbentz and Stanley H. Moore.
January 14, 1948. 36p. diagrs. (NACA
RM E7H27)

Presents development and experimental verification
of analytical method that gives two independent
means of obtaining total-temperature ratio across a
ram jet or turbojet tail-pipe burner without direct
measurement of final gas temperature. Proposes
several total-temperature-ratio meter designs,
which provide simple means of evaluating total-
temperature ratio and also provide basis for ram-jet
fuel-metering controls using total-temperature ratio
as control variable.

NACA RM E8D30

VIBRATION SURVEY OF NACA 24-INCH SUPER-
SONIC AXIAL-FLOW COMPRESSOR. Andre J.
Meyer, Jr. and Morgan P. Hanson. July 30, 1948.
45p. diagrs., photos. (NACA RM E8D30)


Vibration investigations were made of two blade cas-
cades in wind tunnels and of blades operated in the
NACA 24-inch supersonic compressor. The results
showed that the blade vibrations were present at all
tunnel and compressor air velocities and were in-
fluenced primarily by tunnel design, simulated cen-
trifugal loading, surging angle of attack, and possi-
ble critical Mach numbers.

NACA RM E50I29a

INTERPRETATION OF BOUNDARY-LAYER
PRESSURE-RAKE DATA IN FLOW WITH A DE-
TACHED SHOCK. Roger W. Luldens and Robert T.
Madden. December 22, 1950. 14p. diagrs.,
photos. (NACA RM E50I29a)


A procedure is presented for determining boundary-
layer quantities from pressure-rake data, which in-
cludes the combined effects of viscous and shock
losses. The analysis utilizes schlieren photographs
of the shock configuration, the continuity of mass
relationship, and the characteristic of the turbulent
boundary layer that its outer edge is defined by a
rapid change in the slope of the Mach number profile.


NACA RM L7A08

EFFECT OF NUMBER OF FINS ON THE DRAG OF
A POINTED BODY OF REVOLUTION AT LOW
SUPERSONIC VELOCITIES. N. Mastrocola.
April 7, 1947. lOp. diagrs., photos. (NACA
RM L7A08)

The interference drag increased with Increased
number of fins up to a Mach number of 1.35; above
this value the effect is reversed. The magnitude of
interference effects, for the bluff fin sections used
in these tests, is such as to make these effects im-
portant in estimating the drag of a multifin tail
group. The fin drag was found to be comparatively
large and was attributed to the blunt leading edge
and square trailing edge of the fin airfoil section.







26

The test data from flight tests of three- and five-fin
bodies up to a Mach number of approximately 1.4 are
given.



NACARM L7B20

COMPARATIVE DRAG MEASUREMENTS AT TRAN-
SONIC SPEEDS OF 6-PERCENT-THICK AIRFOILS
OF SYMMETRICAL DOUBLE-WEDGE AND
CIRCULAR-ARC SECTIONS FROM TESTS BY THE
NACA WING-FLOW METHOD. Norman S. Silsby.
April 8, 1947. lOp. diagrs. (NACA RM L7B20)


Both airfoils had a thickness of 6 percent of the
chord, were of rectangular plan form, and had an
aspect ratio of 4.0. The tests were run at M = 0.65
to 1.10. The results indicated that the principal
difference in the drag characteristics of the two air-
foils at zero lift is the earlier drag rise of the
double-wedge section. Although the double-wedge
airfoil had a somewhat higher drag throughout the
Mach number range tested, the difference decreased
with increasing Mach number after the onset of the
drag rise of the circular-arc section, and at
M = 1.10 the drag coefficient for the two airfoils
was about the same.



NACA RM L7C05

DRAG CHARACTERISTICS OF RECTANGULAR AND
SWEPT-BACK NACA 65-009 AIRFOILS HAVING
VARIOUS ASPECT RATIOS AS DETERMINED BY
FLIGHT TESTS AT SUPERSONIC SPEEDS. Warren
A. Tucker and Robert L. Nelson. April 22, 1947.
15p. diagrs., photos. (NACA RM L7C05)


Wings of aspect ratios 3.8 to 5.0 were used at Mach
numbers 1.0 to 1.3. The drag coefficient decreased
as the sweepback angle increased, the rate of de-
crease being somewhat greater for the larger aspect
ratios. For Mach numbers greater than a value
somewhat less than that at which the Mach line lies
along the leading edge, the drag coefficient de-
creased with a decrease in aspect ratio.




NACA RM L7K12

PRESSURE DISTRIBUTION OVER A SHARP-NOSE
BODY OF REVOLUTION AT TRANSONIC SPEEDS
BY THE NACA WING-FLOW METHOD. Edward
C. B. Danforth and J. Ford Johnston. March 5,
1948. 25p. diagrs., photos. (NACA RM L7K12)


Contains pressure distributions by the NACA wing-
flow method at Mach numbers between 0.70 and 1.05,
over a sting-mounted body of revolution of circular-
arc profile and fineness ratio 6. The measurements
are compared with existing subsonic and supersonic
theory where applicable. The manner in which
transition occurs from the subsonic to the supersonic
type of pressure distribution is shown with particular
reference to the accompanying rapid increases in
pressure drag.


N


NACA
RESEARCH ABSTRACTS NO. 57

NACA RM L8F23

A LIMIT PRESSURE COEFFICIENT AND AN ESTI-
MATION OF LIMIT FORCES ON AIRFOILS AT
SUPERSONIC SPEEDS. John P. Mayer.
August 23. 1948. 18p. diagrs. (NACA RM L8F23)


The results of an estimation of the limit forces on
airfoils at supersonic speeds are presented. Com-
puted values of the forces on two-dimensional wings
are in good agreement with three-dimensional wind-
tunnel data at high angles of attack where detached
shock waves are present. A limit pressure coeffi-
cient attainable on an airfoil is given as based on
experimental data. The empirical limit pressure
coefficient corresponds to about 70 percent of the
pressure coefficient for a vacuum.



NACA RM L8H06

PRESSURE DISTRIBUTIONS OVER A WING-
FUSELAGE MODEL AT MACH NUMBERS OF 0.4
TO 0.99 AND AT 1.2. Clarence W. Matthews.
November 3, 1948. 24p. diagrs. (NACA
RM L8H06)

Pressure coefficients and Mach numbers are pre-
sented for the flow over a prolate spheroid and an
NACA 65-010 wing section in the transonic Mach
number range. The values over the prolate spheroid
are compared with theoretical values. A study Is
made of the development of a supersonic flow pattern.
Tunnel-wall-constriction effects at Mach numbers
near unity are also considered.



NACA RM L8H3la

EFFECT OF STRUT-MOUNTED WING TANKS ON
THE DRAG OF NACA RM-2 TEST VEHICLES IN
FLIGHT AT TRANSONIC SPEEDS. Sidney R.
Alexander. November 18, 1948. 13p. diagrs.,
photos. (NACA RM L8H31a)


Results of a free-flight investigation near zero lift
of an NACA RM-2 drag research model equipped
with strut-mounted wing tanks of fineness ratio 7.44
are presented for a Mach number range from about
0.7 to 1.1. The addition of the tanks and struts
caused the drag rise to occur at a lower Mach num-
ber and produced a drag-coefficient Increment of
0.075 at a Mach number of 0.72 which Increased to
0.82 (the maximum increment obtained) at a Mach
number of 1.06. Estimation Indicates that the prox-
imity of the struts and tanks may produce significant
trim changes in the Mach number range investigated.




NACA RM L8L07a

FLIGHT INVESTIGATIONS AT LOW SUPERSONIC
SPEEDS TO DETERMINE THE EFFECTIVENESS
OF CONES AND A WEDGE IN REDUCING THE
DRAG OF ROUND-NOSE BODIES AND AIRFOILS.
Sidney R. Alexander. March 3, 1949. 15p. diagre.,
photos. (NACA RM L8L07a)








NACA
RESEARCH ABSTRACTS NO. 57

It is clearly indicated from results and tests con-
ducted at low supersonic speeds that a small cone
placed ahead of a round-nose body can effectively
reduce the drag of the basic body. The presence of
a small wedge placed ahead of a round-nose airfoil
did not appreciably affect the drag of the basic airfoil
in the investigated Mach number range of M = 1.05
to 1.225.



NACA RM L9AI2

AERODYNAMIC CHARACTERISTICS OF A 6-
PERCENT-THICK SYMMETRICAL DOUBLE-WEDGE
AIRFOIL AT TRANSONIC SPEEDS FROM TESTS BY
THE NACA WING-FLOW METHOD. Lindsay J.
Lina. March 4, 1949. 27p. diagrs. (NACA
RM L9A12)

Tests were made in the transonic speed range by the
NACA wing-flow method to investigate the lift,
pitching-moment, and drag characteristics of a 6-
percent-thick-chord symmetrical double-wedge air-
foil having a rectangular plan form of aspect ratio
4.0. The tests covered a range of Mach numbers
from 0.66 to 1.12.


NACA RM L9B17

EXPERIMENTAL DETERMINATION OF THE SUB-
SONIC PERFORMANCE OF A RAM-JET UNIT CON-
TAINING THIN-PLATE BURNERS. John R. Henry.
June 29, 1949. 54p. diagrs., photos. (NACA
RM L9B17)

A ram-jet unit containing a cluster of thin-plate
burners in a semicircular combustion chamber was
tested in Langley induction aerodynamics laboratory.
Data were taken over a fuel-air-ratio range from 0
to 0.049, at combustion-chamber inlet velocities
from 40 to 195 feet per second, and at simulated
free-stream Mach numbers from 0.20 to 0.55.
Combustion efficiencies from 56 to 72 percent were
obtained. Combustion-chamber characteristics led
to the conclusion that high-thrust-output operation
would not be feasible. Thrust-coefficient estimates
for supersonic flight are regarded as too low to be
practical. Diffuser efficiencies of 99 percent were
obtained.


NACA RM L9F02

FLIGHT INVESTIGATION AT HIGH-SUBSONIC,
TRANSONIC, AND SUPERSONIC SPEEDS TO DE-
TERMINE ZERO-LIFT DRAG OF BODIES OF
REVOLUTION HAVING FINENESS RATIO OF 6.04
AND VARYING POSITIONS OF MAXIMUM DIAM-
ETER. Ellis R. Katz. August 31, 1949. 17p.
diagrs., photo. (NACA RM L9F02)


Flight investigation of rocket-powered models was
performed at high-subsonic, transonic, and super-
sonic speeds to determine the zero-lift drag of fin-
stabilized bodies of revolution differing only in
position of maximum diameter. The bodies were of
6.04 fineness ratio and had cut-off sterns with equal
base area for all models.


27

NACA RM L91108

AN EMPIRICAL METHOD FOR ESTIMATING
TRAILING-EDGE LOADS AT TRANSONIC SPEEDS.
T. H. Skopinskld. October 6, 1949. 43p. diagrs.,
tab. (NACA RM L9H08)


An empirical method for estimating trailing-edge
loads and moments in the transonic region is pre-
sented. A comparison of the experimental trailing-
edge normal-force and bending-moment coefficients
obtained for 15 different NACA airfoil sections with
those calculated by the proposed empirical method
indicated that the method may be used as a guide
when specific applicable experimental data are not
available at the design stage.



NACA RM L9J11

PRELIMINARY INVESTIGATION OF A VARIABLE
MASS-FLOW SUPERSONIC NOSE INLET. Clyde
Hayes. December 13, 1949. 15p. diagrs., photos.
(NACA RM L9J11)


A method employing an Inflatable boot on the central-
body surface has been analyzed for varying the mass
flow of supersonic inlets having a circular cross
section and a central body. The tests made at a
Mach number of 2.70 show effects of such a method
on the entering flow, mass flow, and pressure
recovery.



NACA RM L52E29a

PRELIMINARY INVESTIGATION OF THE EFFECTS
OF HEAT TRANSFER ON BOUNDARY-LAYER
TRANSITION ON A PARABOLIC BODY OF REVOLU-
TION (NACA RM-10) AT A MACH NUMBER OF 1.61.
K R. Czarnecki and Archibald R. Sinclair July
1952. 23p diagrs photos., tab. (NACA
RM L52E29a)

This paper presents the results of a preliminary in-
vestigation of the effects of heat transfer on boundary-
layer transition on a parabolic body of revolution
(NACA RM-10) at Mach number of 1.61 This paper
includes also a study of the effectiveness of cooling
on boundary-layer transition with model surface
roughened and a comparison of the results obtained
in this investigation with other available theoretical
and experimental data




NACA RM L53B25

AN EXTENSION OF THE INVESTIGATION OF THE
EFFECTS OF HEAT TRANSFER ON BOUNDARY-
LAYER TRANSITION ON A PARABOLIC BODY OF
REVOLUTION (NACA RM-10) AT A MACH NUMBER
OF 1.61. K. R. Czarnecki and Archibald R.
Sinclair. March 1953. 21p. diagrs., photo.
(NACA RM L53B25)








28


This paper covers the extension of a previous inves-
tigation of the effects of heat transfer on boundary-
layer transition to higher Reynolds numbers, to
greater amounts of heating, and to a more extensive
study of the effects of surface roughness and wind-
tunnel flow disturbances. The tests were made at a
Mach number of 1. 6 and over a Reynolds number
range from 2. 5 x 106 to 35 x 106. A comparison is
made between the experimental results and theory.




THE FOLLOWING REPORT
HAS BEEN DECLASSIFIED FROM
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RM E50KO2

THE FOLLOWING REPORTS
HAVE BEEN DECLASSIFIED FROM
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RM E8E12
RM E8F14
RM E8H06

THE FOLLOWING REPORT
HAS BEEN DECLASSIFIED FROM
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N
0%

NACA
RESEARCH ABSTRACTS NO. 57


UNIVERSITY OF FLORIDA


312620815 02


NACA-Langley 1-39-54 4M




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