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 )
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serial   ( sobekcm )
federal government publication   ( marcgt )
abstract or summary   ( marcgt )

Notes

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

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 001469326
notis - AGY1019
oclc - 01471285
lccn - 86657025
issn - 0499-9274
Classification:
lcc - TL501 .U5895
System ID:
AA00009235:00087

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onal Advisory Committee for Aeronautics


Research Abstracts


N0.26


JULY 23, 1952


CURRENT NACA REPORTS


NACA TN 2711

THE AERODYNAMIC DESIGN OF HIGH MACH NUM-
BER NOZZLES UTILIZING AXISYMMETRIC FLOW
WITH APPLICATION TO A NOZZLE OF SQUARE
TEST SECTION. Ivan E. Beckwith, Herbert W.
Ridyard and Nancy Cromer. June 1952. 30p.
diagrs., 5 tabs. (NACA TN 2711)

A method is given for the design of three-dimensional
nozzles utilizing axisymmetric flow to produce uni-
form flow in a test chamber of arbitrary cross sec-
tion. The method is applied to obtain the coordinates
of a Mach number 10 nozzle for which a square test
section is specified. Radial flow is used in a portion
of the flow field to reduce the computation time. The
remainder of the flow field is computed by the method
of characteristics, but a simplified method is used
near the axis. Tables which facilitate computation of
the radial flow and the flow near the axis are included.
Transition streamlines determined from the analytic
expressions of Kuno Foelsch are compared with the
streamlines obtained from the characteristic net of
the Mach number 10 nozzle.


NACA TN 2720

EFFECTS OF ASPECT RATIO ON AIR FLOW AT
HIGH SUBSONIC MACH NUMBERS. W. F. Lindsey
and Milton D. Humphreys. July 1952. 10p. photos.,
diagrs. (NACA TN 2720. Formerly NACA
RM L8G23)

Schlieren photographs were used in an investigation to
determine the effects of changing the aspect ratio from
infinity to 2 on the air flow past a wing at high sub-
sonic Mach numbers. The results indicated that the
decreased effects of compressibility on drag coeffi-
cients for the finite wing are produced by a reduction
in the compression shock and flow separation.


NACA TN 2724

TRANSONIC SIMILARITY RULES FOR LIFTING
WINGS. Keith C. Harder. June 1952. 26p. diagr.
(NACA TN 2724)

Transonic similarity rules for lifting wings at small
angles of attack are derived. These rules make it
possible to present each aerodynamic coefficient for
a family of wings of varying aspect ratio and thick-
ness ratio in a single design chart for the transonic
range.


NACA TN 2726

ON THE APPLICATION OF TRANSONIC SIMILAR-
ITY RULES. John R. Spreiter. June 1952. 45p.
diagrs. (NACA TN 2726)

The transonic aerodynamic characteristics of wings
of finite span are discussed from the point of view of
a unified small perturbation theory for subsonic,
transonic, and supersonic flows about thin wings.
This approach avoids certain ambiguities which ap-
pear if one studies transonic flows by means of equa-
tions derived under the more restrictive assumption
that the local velocities are everywhere close to
sonic velocity. The relation between the two methods
of analysis of transonic flow is examined, the simi-
larity rules and known solutions of transonic flow
theory are reviewed, and the asymptotic behavior of
the lift, drag, and pitching-moment characteristics of
wings of large and small aspect ratio is discussed.
It is shown that certain methods of data presentation
are advantageous for the effective display of these
characteristics.


NACA TN 2727

EXPERIMENTS IN EXTERNAL NOISE REDUCTION
OF A SMALL PUSHER-TYPE AMPHIBIAN AIR-
PLANE. John P. Roberts and Leo L. Beranek,
Aeronautical Research Foundation. July 1952. 142p.
diagrs., photos., 3 tabs. (NACA TN 2727)

As part of a program to find practicable ways of re-
ducing the external noise level of light airplanes,
noise measurements were made on a representative
pusher-type amphibian airplane. Tests were made
with a standard airplane and with modified versions
of the airplane equipped with a geared engine, exhaust
muffler, and propellers with various numbers of
blades. Sound-level recordings were made of take-
offs; of overhead flights at 100- and 500-foot altitude;
and, for some configurations, of flights at 500-foot
altitude passing 3000 feet away. The tests also in-
cluded analyses of sound-frequency components with
the airplane on the ground from a distance of 50 feet
and at various positions around the airplane. The
results are compared with measurements from a
previous noise study of a tractor-type airplane.


NACA TN 2730

CHOKING OF A SUBSONIC INDUCTION TUNNEL BY
THE FLOW FROM AN INDUCTION NOZZLE. W. F.
Lindsey. July 1952. 20p. diagrs. (NACA TN 2730)


*AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASHINGTON 25, D. C, CITING CODE NUMBER ABOVE EACH TITLE;
THE REPORT TITLE AND AUTHOR.







2


A decrease in Mach number with increase in induc-
tion jet pressure has been shown experimentally, for
one type of induction tunnel, to be associated with
choking of the induced flow. The analysis of the flow
conditions provides a means of predicting the occur-
rence of induction choking so that it can be adequately
considered in design.


NACA TN 2732

THEORETICAL INVESTIGATION OF VELOCITY
DIAGRAMS OF A SINGLE-STAGE TURBINE FOR A
TURBOJET ENGINE AT MAXIMUM THRUST PER
SQUARE FOOT TURBINE FRONTAL AREA.
(Revised). Leo Cohen. June 1952. 34p. diagrs.,
tab. (NACA TN 2732)

The aerodynamic requirements of a single-stage tur-
bine for a turbojet engine at maximum thrust per
square foot turbine frontal area are presented in the
form of velocity diagrams. An impulse diagram with
supersonic velocities relative to the leading and trail-
ing edges of the rotor is required for practically all
blade speeds and turbine-inlet temperatures investi-
gated.


NACA TN 2733

METHOD FOR CALCULATION OF HEAT TRANSFER
IN LAMINAR REGION OF AIR FLOW AROUND CYL-
INDERS OF ARBITRARY CROSS SECTION (INCLUD-
ING LARGE TEMPERATURE DIFFERENCES AND
TRANSPIRATION COOLING). E. R. G. Eckert and
John N. B. Livingood. June 1952. 71p. diagrs.
(NACA TN 2733)

A method which permits approximation of local heat-
transfer coefficients in the laminar-flow region
around cylinders of arbitrary cross section from
those for wedge-type profiles is extended to include
the effects of large temperature differences and tran-
spiration cooling. Charts prepared from exact solu-
tions of the laminar boundary-layer equations for
wedge-type profiles which allow for these effects yield
results with a minimum of calculation. Application
of the method to circular and elliptic transpiration-
cooled cylinders is made to determine local heat-
transfer coefficients and surface temperatures and to
determine the variation in coolant flow required for
maintaining a constant surface temperature.


NACA TN 2735

AN ANALYSIS OF THE NORMAL ACCELERATIONS
AND AIRSPEEDS OF A TWO-ENGINE TYPE OF
TRANSPORT AIRPLANE IN COMMERCIAL OPERA-
TIONS ON ROUTES IN THE CENTRAL UNITED
STATES FROM 1948 TO 1950. Walter G. Walker
and Paul W. J. Schumacher. July 1952. 30p.
diagrs., 4 tabs. (NACA TN 2735)

Normal-acceleration and airspeed data obtained from
a two-engine type of transport airplane are analyzed
to determine the gust and gust-load experiences of
the airplane. The gust and gust-load experiences are
compared with similar results from the same type of


NACA
RESEARCH ABSTRACTS NO.26


airplane during past operations. The influences of
the gusts and the operating speeds in rough air on the
gust-load experience are indicated.


NACA TN 2736

TWO-DIMENSIONAL SHEAR FLOW IN A 90 ELBOW.
James J. Kramer and John D. Stanitz. July 1952.
44p. diagrs. (NACA TN 2736)

As part of an approach to a better understanding of
the motion of real fluids in flow machinery, two-
dimensional, incompressible, nonviscous shear flows
in a 90 elbow have been investigated. Solutions are
presented for linear and sinusoidal velocity distribu-
tions across the inlet of the elbow. The solutions
with linear inlet velocity distributions indicate that as
the negative vorticity of the flow increased: (1) the
static-pressure drop through the elbow decreased,
(2) the local deceleration along the outer channel
walls increased, and (3) the magnitude of the veloci-
ties on the channel walls changed greatly, but the
local pressure coefficient rose only gradually and the
difference in pressure coefficient at corresponding
points on the two walls was practically unchanged. In
the case of a sinusoidal inlet velocity distribution,
local decelerations occurred on both walls.


NACA TN 2738

A PROBABILITY ANALYSIS OF THE METEORO-
LOGICAL FACTORS CONDUCTIVE TO AIRCRAFT
ICING IN THE UNITED STATES. William Lewis and
Norman R. Bergrun. June 1952. 93p. diagrs.,
11 tabs. (NACA TN 2738)

Meteorological icing data obtained in flight in the
United States are analyzed statistically and methods
are developed for the determination of (1) the various
simultaneous combinations of the three basic icing
parameters (liquid-water content, drop diameter, and
temperature) which would have equal probability of
being exceeded in flight in any random icing encoun-
ter, and (2) the probability of exceeding any specified
group of values of liquid-water content associated
simultaneously with temperature and drop-diameter
values lying within specified ranges.


NACA TN 2740

EXPERIMENTAL INVESTIGATION OF THE LOCAL
AND AVERAGE SKIN FRICTION IN THE LAMINAR
BOUNDARY LAYER ON A FLAT PLATE AT A MACH
NUMBER OF 2.4. Randall C. Maydew and Constantine
C. Pappas. July 1952. 22p. diagrs. (NACA
TN 2740)

Average and local skin-friction coefficients for lami-
nar flow have been determined experimentally on a
flat plate at a Mach number of 2.4 for a Reynolds num-
ber range of 0.72 x 106 to 2.8 x 106 and compared
with the laminar-boundary-layer theory of Chapman
and Rubesin. The average skin-friction data were
considerably higher than values predicted by laminar-
boundary-layer theory. Local skin-friction crJeffi-
cients, evaluated by the two methods, agreed well
with the theory when plotted against a momentum-
thickness Reynolds number.


^1







NACA
RESEARCH ABSTRACTS NO.26


NACA TN 2742

BOUNDARY-LAYER DEVELOPMENT AND SKIN
FRICTION AT MACH NUMBER 3.05. Paul F.
Brinich and Nick S. Diaconis. July 1952. 49p.
diagrs., photos. (NACA TN 2742)

Boundary-layer studies consisting of schlieren obser-
vations and momentum surveys were made on hollow
cylinder models with their axes alined parallel to the
stream. Results were obtained for three model diam-
eters and for natural and artificially induced turbulent
boundary-layer flows. Transition Reynolds numbers
were found to decrease with decreases in leading-
edge thickness and with reductions in tunnel pressure
level. Turbulent temperature-recovery factors gen-
erally decreased with increasing Reynolds number
and were a maximum for the smallest transition
Reynolds numbers. The results of this investigation
appeared to be consistent with the theoretical turbu-
lent friction formulas of Wilson and with the extended
Frankl-Voishel analysis of Rubesin, Maydew, and
Varga. Velocity profiles in the outer portion of the
boundary layer could be approximated reasonably with
a 1/7 power profile and were found to be approxi-
mately similar in this region. Velocity profiles given
by the Krmin universal turbulent boundary-layer
Profile parameters were found to be similar in the
laminar sublayer and in the turbulent region.


NACA RM E52D17

EXPERIMENTAL INVESTIGATION OF AVERAGE
HEAT-TRANSFER AND FRICTION COEFFICIENTS
FOR AIR FLOWING IN CIRCULAR TUBES HAVING
SQUARE-THREAD-TYPE ROUGHNESS. Eldon W.
Sams. June 1952. 43p. diagrs., photos. (NACA
RM E52D17)

A heat-transfer investigation was conducted with air
flowing through electrically heated Inconel tubes
having various degrees of square-thread-type rough-
ness (conventional roughness ratios of 0 (smooth tube),
0.016, 0.025, and 0.037), an inside diameter of 1/2
inch, and a length of 24 inches over ranges of bulk
Reynolds number up to 350,000 and average inside-
tube-wall temperatures up to 1950 R. The data
showed that both heat transfer and friction increased
with increase in surface roughness (particularly at
the higher Reynolds numbers) and were also influ-
enced by tube wall-to-bulk temperature ratio. Good
correlation of the heat-transfer data for all tubes was
obtained by use of a modified Reynolds number includ-
ing the friction velocity. Correlation of the friction
data for the rough tubes was obtained for the complete
turbulence region by use of a modified correlation in-
corporating a roughness parameter which excludes
the conventional roughness ratio; no over-all corre-
lation was obtained for the region of incomplete
turbulence.


NACA RM E52D24

SOLUBILITY OF WATER IN HYDROCARBONS. R. R.
Hibbard and R. L. Schalla. July 1952. 25p. diagrs.,
3 tabs. (NACA RM E52D24)


3


An equation is presented which may be used to esti-
mate the solubility of water in any nonolefinic hydro-
carbon or hydrocarbon blend at any temperature.
Only the temperature and the hydrogen-to-carbon
ratio of the hydrocarbon need be known to calculate
solubilities which are believed to be sufficiently
accurate for most engineering purposes.

NACA RM E52E19

USE OF FENCES TO INCREASE UNIFORMITY OF
BOUNDARY LAYER ON SIDE WALLS OF SUPER-
SONIC WIND TUNNELS. Rudolph C. Haefeli. July
1952. 15p. diagrs., photos., tab. (NACA
RM E52E19)

An investigation of the use of solid fences installed
on the side walls of a supersonic wind tunnel to re-
tard the development of transverse flow and thus to
increase the uniformity of the side-wall boundary
layer is reported. Beneficial results were obtained
with fences which had depths of the order of the
boundary-layer displacement thickness and which
followed potential-flow streamlines through the
nozzle. Reduction of the number of fences on each
side wall from four to two eliminated their effective-
ness.


NACA TM 1223

SOME EXPERIENCES REGARDING THE NONLINE-
ARITY OF HOT WIRES. (Quelques experiences sur
la non-line'arit6 des fils chauds). R. Betchov and
W. Welling. June 1952. 13p. diagrs. (NACA
TM 1223. Trans. from Koninklijke Nederlandsche
Akademie van Wetenschappen te Amsterdam, Ver-
handelingen, v.53, no.4, 1950, p.432-439; Technische
Hoogenschool te Delft, Laboratorium voor Aero- en
Hydrodynamica. Mededeling 66).

Hot-wire experiments illustrate some differences be-
tween the resistance characteristics as given by
King's equation and those for a finite wire length.
The experiments support previous theoretical work
which accounted for a heat loss although the wire ends
and a dependence of the convective heat loss on a non-
linear function of the temperature difference between
the wire and the air.


NACA TM 1346

NONLINEAR THEORY OF A HOT-WIRE ANEMOM-
ETER. (Theorie non-linealre de l'anemometre a fil
chaud). R. Betchov. July 1952. 23p. diagrs.
(NACA TM 1346. Trans. from Koninklijke Neder-
landsche Akademie van Wetenschappen te Amsterdam,
Verhandelingen, v. 52, no. 3, 1949, p. 195-207;
Technische Hoogenschool te Delft, Laboratorium
voor Aero- en Hydrodynamica. Mededeeling 61).

A theoretical analysis is presented for the hot-wire
anemometer to determine the differences in resist-
ance characteristics as given by King's equation for
an infinite wire length and those given by the addi-
tional considerations of (a) a finite length of wire
with heat loss through its ends and (b) heat loss due
to a nonlinear function of the temperature difference
between the wire and the air.








4



BRITISH REPORTS


N-6183

Ministry of Supply (Gt. Brit.)
THE MECHANICAL CHARACTERISTICS OF
PNEUMATIC TYRES. R. Hadekel. March 1960.
iii, 146p. diagrs., photos. (MOS S RESTRICTED

The object of this report is to give an exposition of
the existing kii-wlJ.dge- of the mechanical properties.
characteristics and behavior of pneumatic tires, in-
sofar as they may be of interest to the user. Ques-
tions of tire terhn,,h'-y are not discussed. The roll-
ing i|r_.,:ers. vertical and horizontal load-deflection
characteristics, behavior during braking and side-
ways motion, rolling resistance and high speed ef-
fects are all dealt with. Where possible, both
theoretical treatment and empirical data are given.
Recommendations are made with respect to further
research. The primary object of the investigation
was to provide data for aircraft purposes, but it is
hoped that the information here presented will also be
of use to land vehicle designers.


N-14407"

Aeronautical Research Council (Gt. Brit.)
SANDWICH CONSTRUCTION AND CORE MATERIALS.
PART V. W. J. Pullen. 1952. 14p. diagrs.,
photos., 5 tabs. (ARC R : M 2686; ARC 10,501.
Formerly ARC 9017; Strut 938; Plas. 57)

Tensile, compressive and creep tests have been
carried out on four different samples of extruded
cellular cellulose acetate. It is concluded that the
material is comparable with calcium alginate and
other low-density materials so far handled in the
Engineering Division, N.P.L. In particular, the
samples are not subject to the same degree of
"softening" as has been the case with some similar
materials. The "filled" samples are more efficient
than the "unfilled" ones and are worth considering as
possible low-density stabilizers in sandwich construc-
tion.


N-14772*

Aeronautical Research Council (Gt. Brit.)
A REVIEW OF SOME STALLING RESEARCH. A. D.
Young. WITH AN APPENDIX ON WING SECTIONS
AND THEIR STALLING CHARACTERISTICS. H. B.
Squire and A. D. Young. 1951. 39p. diagrs., 5 tabs.
(ARC R & M 2609; ARC 5751. Formerly RAE
Aero 1718)

Over a period of years a considerable amount of stall-
ing research on various airplanes was completed at
the Royal Aircraft Establishment and it was consid-
ered desirable that the main results should be sum-
marized and reviewed. The report includes a general
discussion of the effect on stalling behavior of wing
section, plan form, washout, flaps, nacelles, gills,
slipstream, automatic wing-tip slots and Hudson-


NACA
RESEARCH ABSTRACTS NO.26


type slits. The important part that is played by the
longitudinal trim and stability at incidences near the
stall is emphasized. The relation between wing sec-
tions and their stalling characteristics is discussed
and it is shown that the stalling characteristics can be
broadly predicted from an examination of the form of
the wing-section upper-surface pressure distribution
at high incidences. The results indicate that vicious
stalling behavior can be avoided by the use of wing
sections towards the tip of fairly high camber (3 to
4 percent) and moderate thickness (> 12 percent).
For some types of airplanes there are, however,
serious objections to the use of high camber towards
the tips; the designer is then advised to avoid wing
sections which experiments and theory indicate have
particularly bad stalling characteristics. The worst
tip thickness for stalling appears to be in the region
of 9 percent. High taper tends to worsen the stalling
behavior and it is advisable to consider taper ratios
greater than 2:1 only in conjunction with wing-tip
sections having good stalling characteristics. The
use of part-span flaps does not appear to cause any
marked deterioration in stalling behavior, and fre-
quently it improves the behavior; but there is some
evidence, though not yet conclusive, that the use of
full-span flaps may be accompanied by an appreciable
worsening in stalling behavior. Attention is drawn to
the advisability of examining the flow at high inci-
dences in the neighborhood of the tail-plane of an air-
plane in the design stage, with a view to assessing its
probable stalling behavior; In particular, the possi-
bilities of designing for some stall warning can then be
examined.


N-16033'

Aeronautical Research Council (Gt. Brit.)
THE DIFFUSION OF TRANSVERSE LOADS IN A
REINFORCED CIRCULAR CYLINDER WITH NON-
RIGID FRAMES. S. R. Lewis. 1952. 23p. diagrs.
(ARC CP 74)

Explores the variation of shear stress distribution at
a loaded frame consequent upon the variation of sev-
eral geometric parameters such as frame stiffness,
skin thickness, stringer spacing, etc. Figures indi-
cate the shear distribution around a frame for a sin-
gle concentrated radial load of 1000 lb. The param-
eters chosen are those common to aircraft design,
and it is possible to obtain a reasonably accurate
shear distribution around a frame from the data sup-
plied, without doing the actual lengthy shear calcula-
tion. The work is confined to the particular case
where the loaded frame is removed from the end of
the cylinder. Gives method of obtaining the shear
load due to a tangential load and moment from the
radial load expressions.


N-16034*

Aeronautical Research Council (Gt. Brit.)
DIGITAL RECORDING AND ANALYSING OF FLIGHT
TEST DATA: A PROPOSED SYSTEM. E. J.
Petherick. 1952. 0lp. diagrs. (ARC CP 75)

A small digital recorder is proposed, to punch 10,000
instrument readings on 100 feet of cine film, each







NACA
RESEARCH ABSTRACTS NO.26


item to 3 decimal or 12 binary places. Any number
of instruments could be recorded at 10 readings a
second. A further unit is envisaged, to read the
punched data; to correct each item for instrument
errors; to display the corrected values; and to punch
them on Hollerith cards or RAESCC tape.


N-16036"*

Aeronautical Research Council (Gt. Brit.)
THE EFFECTS OF ATMOSPHERIC HUMIDITY AND
TEMPERATURE ON THE ENGINE POWER AND
TAKE-OFF PERFORMANCE OF A HASTINGS I.
G. Jackson. 1952. 34p. diagrs., 9 tabs. (ARC
CP 77)

Engine power decreases with increasing humidity and
reduction is greatest at take-off engine speed and
boost. Specific humidity will rarely exceed 2-1/2
percent and this will decrease take-off power approxi-
mately 10 percent as compared with dry air. The
take-off distance to clear a 50 ft. obstacle would be
increased about 17 percent. For take-off conditions,
humidity is a parameter of the same order of impor-
tance as temperature. At constant humidity the rate
of decrease of power with increase of temperature is
not very different from the value given by the stand-
ard formula below full-throttle height. No effect of
humidity on fuel consumption has been detected. Fuel
consumption decreases with increase in temperature.



N-16038*

Aeronautical Research Council (Gt. Brit.)
AN EXPERIMENTAL INVESTIGATION OF THE
PERFORMANCE OF A PILOT PLANT FOR DRYING
AIR BY SOLID GRANULAR ADSORBENTS. P. J.
Bateman. 1952. 44p. diagrs., 2 tabs. (ARC CP 79)

An investigation was undertaken to measure the per-
formance on a pilot plant scale of uncooled beds of
granular adsorbent when drying air to the low humid-
ities required for supersonic wind tunnel operation.
Two size grades of activated alumina and one of silica
gel were tested in a bed 2-1/2 by 2-1/2 feet in plan,
and of various depths between 12 and 25 inches. The
bed was worked throughout at about atmospheric pres-
sure and air was drawn from atmosphere without
treatment for humidity or temperature. The results
are presented, in the main, as graphs of mean bed
concentration mass of water adsorbed
\mass of activated adsorbent) against
inlet humidity for two values of the outlet humidity,
0.0005 and 0.0002 lb water/lb air. It is shown that
over the ranges of the variables encountered the mean
bed concentration decreases with increasing inlet
humidity, increases with increasing bed depth at a
fixed contact time, and decreases with increasing
grain size. Although a definitive comparison is not
possible there does not appear to be much difference
in the performance of silica gel and activated alumina
of the same grain size.


5


N-16039T

Aeronautical Research Council (Gt. Brit.)
CALCULATED PRESSURE DISTRIBUTIONS FOR THE
R.A.E. 100-104 AEROFOIL SECTIONS. R. C.
Pankhurst and H. B. Squire. 1952. 20p. diagrs.,
3 tabs. (ARC CP 80)

Gives the positions of maximum thickness, leading-
edge radii, trailing-edge angles, coordinates, shapes,
pressure distributions, and velocity distributions for
the RAE 100-104 airfoil sections.


N-16042*

Aeronautical Research Council (Gt. Brit.)
WIND TUNNEL TESTS ON A 90 APEX DELTA WING
OF VARIABLE ASPECT RATIO (SWEEPBACK 36.80).
PART I. GENERAL STABILITY. J. G. Ross, R.
Hills and R. C. Lock. PART E. MEASUREMENTS
OF DOWNWASH AND EFFECT OF HIGH LIFT DE-
VICES. R. C. Lock, J. G. Ross and P. Meiklem.
1952. 67p. diagrs., photos., 26 tabs. (ARC CP 83)

Part I. Longitudinal and lateral stability measure-
ments have been made in a low speed tunnel on a delta
wing of 90 apex angle with three different taper
ratios. The tests included measurements with ground,
the effect of a body, and measurements of elevon
power. CLmax was 0.86 for all taper ratios but was
-max
reduced to a trimmed value of 0.65 with a static mar-
gin of 0.10a, due to the large loss of lift caused by the
elevons. A tip stall starts on the wings at a 8 to
12 depending on the taper ratio; this has compara-
tively little effect on pitching moments but a large ef-
fect on both rolling and yawing moments, nv and -zv
both decreasing after the tip stall. C. A. T. tests
suggest that there is an appreciable favorable scale
effect on the tip stall. Ground effects are small and
can be estimated sufficiently accurately using existing
theoretical work on unswept wings. Part 1. Wind-
tunnel measurements of downwash were made on delta
wings of aspect ratios 4, 3, and 2.3, using a tail of
delta plan form in three vertical positions at two
chordwise stations behind the wing. The tests also
included the effect of the tail, and of split flaps and
nose flaps, on the stability near the stall and on
CLmax The tip nose flaps proved effective in delay-
"max f ff
ing the tip stall, and gave some increase in CLmax
max
With split flaps untrimmed CLmax was 0.95 and 1.2
with the flap in the forward and rear position respec-
tively. There was no change of trim with the flaps in
the forward position with tail off; with the tail an in-
termediate flap position should give zero trim change.
The downwash was large at high lift coefficients, ow-
ing to the early tip stall, and this caused a loss of tail
efficiency with a corresponding slight instability near
the stall, which should not be serious. A method is
given of calculating the downwash at small incidences
behind a delta wing, and the results show good agree-
ment with the measured values, provided that the ex-
perimental lift-curve slope is used.








6


N-16047*

Aeronautical Research Council (Gt. Brit.)
A THEORETICAL INVESTIGATION OF THE RE-
SPONSE OF A HIGH-SPEED AEROPLANE TO THE
APPLICATION OF AILERONS AND RUDDERS. K.
Mitchell, A. W. Thorpe and E. M. Frayn. 1952.
92p. diagrs., 6 tabs. (ARC R & M 2294; ARC 8831;
ARC 9639. Formerly RAE Aero 2040; RAE
Aero 1952)

The response of a fast moving airplane to a lateral
gust, and to applied rolling and yawing moments, is
examined by means of the differential analyzer,
taking a range of values of the principal lateral sta-
bility parameters, and including sufficient ranges of
the other stability and inertia parameters to make the
conclusions of general validity for high-speed flight.
The motion following a sharp-edged side-gust is
shown to be a markedly oscillatory character, with
an unpleasantly short period, particularly in small
airplanes. The shortness of the period is probably
the worst feature. A general survey is made of the
dependence of the motion upon the various parameters,
the differential analyzer results being supplemented
by the use of approximate formulas, which were de-
veloped with a view to this application. Particular
attention is paid to the amplitudes of the motion in
roll, yaw, and sideslip, and it is seen that it may be
difficult to make the motion less unpleasant. The
period may be lengthened by reducing nv, but the
improvement that is possible in this way is limited.
Damping can be improved by reducing dihedral, or by
increasing body side area: the addition of a forward
fin, ahead of the center of gravity, would therefore be
doubly helpful, lengthening the period and improving
the damping. In studying response to applied moments
attention is chiefly concentrated upon response to
ailerons, and the theoretical results are compared
with a theoretical standard motion produced by a con-
stant rolling moment together with a yawing moment
varied so as to suppress sideslip. Response at high
speeds is shown to be insensitive to changes in Zv and
nv within their normal ranges, and good response to
pure rolling moment is assured for all lateral sta-
bility characteristics other than those associated with
the combination of small fin with large dihedral: this
combination is worst at high values of the lateral rel-
ative density. The effect of adverse yawing moment
from the ailerons is detrimental, and becomes worse
as the dihedral is increased or fin area decreased.


N-16048*

Aeronautical Research Council (Gt. Brit.)
LATERAL RESPONSE THEORY. K. Mitchell.
APPENDICES. E. M. Frayn. 1952. 57p. 8 tabs.
(ARC R & M 2297; ARC 7933; ARC 8278; ARC 8544.
Formerly RAE Aero 1925; RAE Tech. Note Aero
1513; RAE Tech. Note Aero 1584)

It is usual, when investigating the response of an air-
plane to applied forces and couples, to examine its
behavior when disturbed by rolling and yawing mo-


NACA
RESEARCH ABSTRACTS NO. 26


ments, instantaneously or linearly applied. The
relevant curves can be readily obtained by means of
a differential analyzer: their calculation by existing
methods is extraordinarily tedious, and analysis of
lateral response, on a basis of the curves alone, a
matter of some difficulty. The need was therefore
felt for some new approach to the subject which would
offer an alternative method of assessing response. A
method is now given, whereby the calculation of re-
sponse to any forces whatever is made to depend upon
the prior determination of a number of constants, for
which the name "modal response coefficients"' is pro-
posed. The idea of the method is introduced by solving
first by the method of variation of constants; later
the relevant formulas are more simply obtained by
operational methods. It is shown that the Laplace
transformation scores over Heaviside's methods here
in leading at once to a result of considerable general-
ity and in a form more convenient for computation.
Methods for calculating model response coefficients
and response curves are examined in appendices,
recommended methods being presented as routine
computations to be carried out on printed computing
sheets (given as tables 1 and 2). "Optimum-interval"
tables of certain functions occurring in the calculation
of response curves are also given (tables 3 to 8). As
a method of calculating response curves, even with
the tabular aids presented here, the method of modal
response coefficients is probably slightly less
laborious than the usual methods: this work should be
handled mechanically by the differential analyzer.
The calculation of the modal response coefficients
themselves, however, is not tedious, and may usefully
be undertaken to aid in the study of curves calculated
by other means. It is hoped that, when a little experi-
ence has been gained, it may prove possible to dis-
cuss response in terms of the modal response coeffi-
cients alone.


N-16049*

Aeronautical Research Council (Gt. Brit.)
RELAXATION TIME EFFECTS IN GAS DYNAMICS.
J. C. Gunn. 1952. 44p. diagrs., 12 tabs. (ARC
R & M 2338. Formerly ARC 9536; FM 910; TP 145;
TJR 89)

In the formulation of the equations of gas dynamics it
is usual to make certain simplifying assumptions con-
cerning the properties of the gas under consideration.
Thus, the equations are greatly simplified if it can be
assumed that the gas is a perfect one, satisfying the
equation p/p = RT, and also that the specific heat,
Cp, is a constant. A further assumption, generally
made, is that the energy content of the gas is capable
of instantaneous change following the external condi-
tions, so that the internal energy is always a uniquely
defined function of the temperature. These assump-
tions are, of course, often justified in practice, but
there is need for caution in applying them to unfamil-
iar gases, or, as we shall see, to cases where the
temperature range of the process is very wide. The
object of this note, in particular, is to study the third
of the assumptions referred to above, and we shall in-
vestigate the effect on various dynamical processes of







NACA
RESEARCH ABSTRACTS NO.26


the known time lag in the picking-up, by the vibra-
tional degrees of freedom of a gas, of their equilib-
rium energy. Three main cases will be considered:
(1) flow in a wind tunnel, (2) flow round an airfoil,
without shock waves, and (3) flow through shock
waves. The case of propagation of sound in a medium
exhibiting heat capacity lag is also briefly discussed.
The aim of the consideration is twofold: firstly, to
demonstrate how small the effects of heat capacity lag
are in most aerodynamic phenomena, and secondly to
indicate some special cases in which the effects be-
come more important. The preliminary physical data
concerning the vibrational energy of a gas, and its
adjustment to equilibrium, have been fairly fully pre-
sented in papers by Kantrowitz, and, more particular-
ly by Bethe and Teller2. As these papers may not be
readily accessible the opportunity is first taken to
summarize this basic physical knowledge. Some of
the more general consequences of the heat-capacity
lag are also deduced.


N-16050*

Aeronautical Research Council (Gt. Brit.)
WINGS OF FINITE ASPECT RATIO AT SUPERSONIC
VELOCITIES. D. R. Taunt and G. N. Ward. 1952.
12p. diagrs. (ARC R & M 2421; ARC 9401.
Formerly Department of Scientific Research and
Experiment, Admiralty SRE/Airflow/29)

An error in the well-known work by Schlichting is
corrected, and it is shown that the integral equation
derived can be solved analytically. The main result
is that for a plane rectangular wing of aspect ratio A,
the lift coefficient is less than that for an infinite
plane wing by a factor 1 1 where M is the
2A -2 1
Mach number of the undisturbed stream. It is shown
that the analytical solution can be extended to plane
trapezoidal wings of any cutting-off angle.


N-16051*

Aeronautical Research Council (Gt. Brit.)
THE VALIDITY OF FULL-SCALE PROPELLER
BRAKING TESTS IN THE R. A. E. 24-FT WIND
TUNNEL. A. B. Haines and R. J. Monaghan. 1952.
9p. diagrs., 2 tabs. (ARC R& M 2473; ARC 10,309.
Formerly RAE Tech. Note Aero 1851)

During tests at negative pitch air flow through the
working section is considerably disturbed. To check
the effect of tunnel interference, tests were made in
both the 5-foot and 24-foot wind tunnels. Tests also
provided information on Reynolds number effects and
solidity. It was concluded that the effects of tunnel
interference can be allowed for by a correction to J.
Over the range of reliable measurements this addi-
tivc correction decreases with J, becoming negligible
for J > 1.6. For the tests an increase in Reynolds
number had no appreciable effect on negative thrust.
Correcting for a 2:1 increase in solidity by treating
thrust and power as proportional to solidity may over-
estimate negative thrust for a given power by 30 per-
cent.


7


N-16052*

Aeronautical Research Council (Gt. Brit.)
SUPERSONIC DIFFUSERS. J. Lukasiewicz. 1952.
30p. diagrs., photos., 3 tabs. (ARC R& M 2501;
ARC 10,110. Formerly RAE Gas 8)

Various possible types of supersonic diffuser are
considered theoretically and the available experimen-
tal evidence is reviewed. The most obvious type of
an efficient supersonic diffuser is the reversed super-
sonic nozzle. A stability criterion for the reversed
supersonic nozzle is developed, which shows that the
contraction between the entry and the throat is so
limited that if a normal shock at entry is to be avoided
such a diffuser is of little practical use unless an
artificial means of inducing and maintaining the less
stable flow can be found. Test results of simple pitot
entries show that, in the absence of the boundary
layer, the one-dimensional theory holds for normal
shocks. Since their efficiency is high for Mach num-
ber up to about 1.5, diffusers with normal shock at
entry are quite satisfactory in this velocity range.
The mechanism of the shock wave and boundary-layer
interaction is considered. When normal shocks occur
in the presence of the boundary layer, as in annular
pitot entries and diffusers for supersonic wind tun-
nels, their compression efficiency is appreciably
lower than the theoretical one. For very high veloc-
ities, exceeding M = 2.5, the multishock diffuser is
practicable and offers high theoretical efficiency.
The efficiencies of this type of diffuser are examined
and shown to exceed 90 percent up to M = 3.0 for de-
signs having three or four oblique shocks. Various
geometrical arrangements are considered from the
point of view of reduction of frontal area for a given
mass flow, and improvements on the single focused
wave system are obtained. The high performance of
multishock diffusers is confirmed by Oswatitsch's
tests at Gottingen, the results of which are here
summarized. The operation of supersonic diffusers
at other than the design Mach number is briefly con-
sidered.


N-16053*

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE FRISE AILERON. H. H. B. M.
Thomas and E. R. Crabbe. 1952. 16p. diagrs.
(ARC R & M 2502; ARC 10,213. Formerly RAE
Aero 2163)

In spite of its well known disadvantages the Frise
aileron is still used and can yield a good aileron con-
trol when the balance required is not very close.
There appears to be, therefore, a need for some
means of estimating the hinge-moment characteris-
tics of this type of aileron. The present note, besides
giving a general discussion of the characteristics of
Frise ailerons, includes a method of estimating the
hinge moment from the hinge moment of the unbal-
anced ailerons. Means of improving the hinge mo-
ment characteristics, avoiding oscillation, and in-
creasing the rolling power of Frise ailerons are also
discussed. These include the use of differentially
geared tabs and slats attached to the wing or the
aileron.








8


N-16057*

Aeronautical Research Council (Gt. Brit.)
THE PREVENTION OF BINARY FLUTTER BY
ARTIFICIAL DAMPING. R. A. Frazer. 1951. 16p.
diagrs., 2 tabs. (ARC R & M 2552. Formerly
ARC 7473; ARC 0.397)

Formulas are obtained which provide an estimate of
the amount of artificial control needed to prevent
binary flutter. Results are presented in terms of a
"minimum damping multiplier" defined as the ratio of
the least direct damping coefficient required for
absolute flutter prevention to the "natural" direct
aerodynamic damping coefficient of the control sur-
face concerned. Numerical results are obtained for
five different types of aircraft.



N-16058*

Aeronautical Research Council (Gt. Brit.)
WIND-TULNNEL TESTS ON THE SPOILINu EFFECTS
OF ENGINE COOLING GILLS ON RADIAL AIR-
COOLED INSTALLATIONS ON A WINu J. Seddon
and J. A. Kirk. 1952. 26p. diagrs., 6 tabs. (ARC
R & M 2558; ARC 5659. Formerly RAE Aero 1724)

Spoiling drag associated with fully open gills at high
CL can be very large if the gill exit is nearer to the
wing leading edge than about 10 percent of the local
wing chord; but the effect diminshes rapidly as this
distance is increased. The drag due to spoiling is
reduced if the cooling air is kept away from the
nacelle-wing junction by emitting it at specified re-
gions around the exit, preferably at the bottom where
the lift is a minimum. Larger gill angles would be
needed to satisfy maximum-flow requirements in
this way. The return-flow cooling system, with nose
exit, shows no evidence of large spoiling drag at high
cooling flow. The data obtained might be useful for
estimating effects of other forms of discharge of low-
energy air in front of a wing leading edge.


N-16059*

Aeronautical Research Council (Gt. Brit.)
THE PRODUCTION OF LIFT INDEPENDENTLY OF
INCIDENCE THE THWAITES FLAP, PARTS I
AND II. B. Thwaites. 1952. 20p. diagrs. (ARC
R & M 2611. Formerly ARC 10,100; FM 1020;
Perf. 252; ARC 11,023; FM 1176, Perf. 378)

In Part I of this paper, the possibility of obtaining lift
on a body in a uniform stream independently of the
incidence is discussed, and a practical method which
obtains this effect if given. It is shown that a small
thin "flap" which may be moved about a well-rounded
trailing edge through which, for example, continuous
suction is applied will produce circulation about the
airfoil. A necessary feature of this method is the
prevention of separation of flow by boundary-layer
suction, which is also used to reduce substantially the
width of the wake. The method uses principles quite
different from those which have been proposed in the
past for obtaining increased lift Qn airfoils. The
practical applications of the device are briefly dis-


NACA
RESEARCH ABSTRACTS NO.26


cussed, and some interesting consequences pointed
out. It will, for instance, be possible to fly with an
airfoil always at zero incidence. Again, the stall in
which the flow separates from near the leading edge
may be completely avoided, for as the circulation and
lift increase, the incidence may be decreased so that
severe adverse velocity gradients occur nowhere but
near the trailing edge.

In Part II of the paper, a report is given of a prelimi-
nary experiment which was set up to investigate
whether the theoretical predictions made about the
efficacy of the flap were largely confirmed. A wholly
porous circular cylinder was fitted with the flap and
measurements were made of the pressure distribution
round the cylinder for various positions of the flap.
These observations showed that for angular deflection
of the flap of less than 20, about 85 percent of the
theoretical value of CL was realized: a maximum CL
of about 5.6 was obtained. These results are taken to
show that the physical principles of Part I are sound
and that the Thwaites flap does, in fact, enable lift to
be generated independently of incidence.


N-16060*

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE CHARACTERISTIC CURVE FOR AN
AIRSCREW OR HELICOPTER. C. N. H. Lock. 1952.
3p. diagr. (ARC R & M 2673. Formerly ARC 10,636;
ARC H46)

Modifies the method of plotting the curves as given in
R & M 1026 and R& M 1014. The new variables are
the square roots of the variables used in these earlier
reports. The change of variables has the following
three advantages: (1) The three principal working
states now correspond to three different quadrants,
(2) the representation in the neighborhood of the
x-axis and the y-axis is more definite since the curve
has a finite slope at both these points, and (3) the
formulas of the vortex theory take simple forms.


N-16061*

Aeronautical Research Council (Gt. Brit.)
TANK TESTS ON A HULL WITH THE MAIN STEP
FAIRED IN PLANFORM AND ELEVATION. D. I. T.
P. Llewelyn-Davies. 1952. 36p. diagrs., photos.,
4 tabs. (ARC R & M 2708; ARC 8807. Formerly
RAE Aero 2029)

Tank tests were required to find out whether the water
characteristics of a hull with a main step, faired in
both plan form and elevation, were comparable with
those of a hull with a conventional Vee or transverse
step. Stability diagrams and spray and resistance
characteristics were obtained over a large range of
loadings (CA0 = 0.616 to CA = 1.440). The fully
,\O
faired step offers more possibility of designing a lon-
gitudinally stable flying boat hull than does the conven-
tional transverse or Vee step, but a hull with such a
step is 5 to 10 percent less efficient hydrodynamically
except at high speed. In order to avoid running too
fine at high speed, it is recommended that the center
of gravity should not be more than 0.46b ahead of the








NACA
RESEARCH ABSTRACTS NO.26


apex of the step. The modification to the step plan
form makes little difference to the main spray char-
acteristics, but increase in all-up-weight reduces
wing, tailplaue and propeller clearances. The effect
of increase in load on the porpoising stability char-
acteristics is to raise both limits, with a tendency for
the upper limit to rise more rapidly, but less regular-
ly, than the lower limit. The free-to-trim attitudes
also rise with increase in all-up-weight. The planing
efficiency of the hull increases with increase of load,
especially at high speeds. There is evidence of a
second resistance hump at high speeds and also of a
critical variation of planing efficiency with attitude
under similar conditions.


N-16313"

Royal Aircraft Establishment (Gt. Brit.)
CHARTS OF THE WAVE DRAG OF WINGS AT ZERO
LIFT. T. Lawrence. January 1952. 18p. diagrs.
(RAE Tech. Note Aero 2139)

Theoretical calculations of the wave drag at super-
sonic speeds of nonlifting wings of double wedge and
biconvex section are reviewed, and the best method
of presenting the results considered. Using this
method, a representative selection of the available
numerical evaluations of the theory is presented.
These should be of value for wing drag estimation
purposes.



MISCELLANEOUS



NACA TN 2657

Errata No. 1 on "SOME EFFECTS OF FREQUENCY
ON THE CONTRIBUTION OF A VERTICAL TAIL TO
THE FREE AERODYNAMIC DAMPING OF A MODEL
OSCILLATING IN YAW". John D. Bird, Lewis R.
Fisher and Sadie M. Hubbard. April 1952.


UNPUBLISHED PAPERS


N-8809*

THE PHUGOID OSCILLATION OF THE AIRPLANE
WITH CONSIDERATION OF UNSTEADY AIR FORCES.
(Die Phygoidschwingung des Flugzeugs bei Berick-
sichtigung instationarer Luftkrafte). J. Weissinger.
March 1952. 33p. diagrs., 3 tabs. (Trans. from
Zentrale fur wissenschaftliches Berichtswesen der
Luftfahrtforschung, Berlin. FB 1430; Deutsche
Versuchsanstalt fur Luftfahrt E. V., Berlin. Institute
fuir Aerodynamik, March 29, 1941).

An iteration method for determination of the influence
of the unsteady air forces on the slow longitudinal
oscillation of the airplane (to be calculated according
to the method of small oscillations) is described and
the results of a few numerical calculations are
given.


9


N-9665*

SPECTRAL ENERGY DISTRIBUTION IN RADIATION
DUE TO ENCOUNTERING SHOCK WAVES. (Reparti-
tion spectrale energetique dans la Lumiere emise
lors de la Recontre d'Ondes de Choc). A. Michel-
Levy, H. Muraour and E. Vassy. April 1952. 19p.
diagrs., photos. (Trans. from Revue d'Optique
Theorique et Instrumentale, 1941, p.149-160).

This report deals with the determination of the
energy distribution in a continuous spectrum ob-
tained at the encounter with shock waves in the at-
mosphere of various gases: argon, helium, oxygen,
etc. The reference source is the positive crater of
a carbon arc. Various regions of the gas chamber
were explored. The radiation is not that of a black
body. The present status of the theory of continuous
spectra offers no possibility for interpreting this
phenomenon.


N-12762 *

THE STRESS-STROKE DIAGRAM OF AIRPLANE
SHOCK ABSORBER STRUTS. THIRD PARTIAL
REPORT THE LANDING IMPACT OF OLEO-
PNEUMATIC SHOCK ABSORBERS. (Zur kenntnis der
Kraftwegdiagramme von Flugzeugfederbeinen, 3.
Teilbericht: Der Landestoss von Olluftfederbelnen).
K. Schlaefke. April 1952. 15p. diagrs., 3 tabs.
(Trans. from Aerodynamische Versuchsanstalt
Gottingen E.V., Technische Berichte, v. 11, no. 5,
1944, p. 137-141).

The stress-stroke diagram of an oleo-pneumatic leg
under the first impact at landing can be obtained by
superposing onto the diagram of the undamped com-
pressed air leg, the damping diagram according to
the second approximate method described in the pre-
ceeding partial report. The maximum stroke, max-
imum stress, and efficiency to be expected for divers
damping, lift coefficients, and heights of drop are
computed and represented diagrammatically for prac-
tical use. Lastly, it is indicated how the newly intro-
duced damping factor for oleo-pneumatic legs can be
determined by experiment.


N-12763*

ON THE RECIPROCAL EFFECTS BETWEEN SHOCK
ABSORBER STRUT AND TIRE AT LANDING IMPACT
OF AIRPLANE UNDERCARRIAGES. (Zur Kenntnis
der Wechselwirkungen zwischen Federbein und Reifen
beim Landestoss von Flugzeugfahrewerken). K.
Schlaefke. May 1952. 18p. diagrs., 3 tabs. (Trans.
from Aerodynamische Versuchsanstalt Gottingen
E. V., Technische Berichte, v. 10, no. 11, 1943,
p. 363-367).

This report deals with the reciprocal effects between
shock-absorber strut and tire due to impact at land-
ing. It was found that the tire reaches its maximum
springing earlier than the strut and that at the first
instant of landing impact the energy is taken up solely
by the tire. The results are represented in diagrams
which can be used for the conversion of experimental
data to other combinations of tires and shock-
absorber struts.




UNIVI"SITY OF FLORIDA
II III IIIl 11 I1I11I11III1111l l I1I1 11111l III II IIIIIIII 1
NACA 3 1262 09079 7506
NACA
10 RESEARCH ABSTRACTS NO.26


N-15440*

Massachusetts Inst. of Tech.
ASYMPTOTIC SOLUTIONS OF THE STABILITY
EQUATIONS OF A COMPRESSIBLE FLUID.
Cathleen S. Morawetz. July 1951. 47p. diagrs.
(Massachusetts Inst. of Tech.)

Two conclusions of practical interest are reached
here. It is proved rigorously that the proper branch
of the multiple-valued asymptotic solutions has been
used by Lees and Lin in NACA TN 1115. Establishes
the existence or absence of an inner viscous region
and its nature.


N-15499*

HOT-WIRE ANEMOMETER UTILIZIN, DIRECT-
CURRENT AMPLIFIER. (Chokur Yu Zofukuki No
Riyoseru Nessen-Fusikukeij 8p. diagrs. (Trans.
of Jap. rept., February 26, 1944, Wright-Patterson
Field microfilm, Jap/WTD/Re/336).

Des% ribeb the design of a hot-wire anemometer for
measuring the instantaneous velocity fluctuations by
means of a direct current amplifier and gives the
properties of the instrument as determined experi-
mentally.


DECLASSIFIED NACA REPORTS


NACA RM L8102

INVESTIGATION OF TWO PITOT-STATIC TUBES
AT SUPERSONIC SPEEDS. Lowell E. Hasel and
Donald E. Coletti. November 19, 1948. 24p. diagrs.
(NACA RM L8102) (Declassified from Confidential,
6/4/52)

The results of tests at a Mach number of 1.94 of an
ogival-nose cylindrical pitot-static tube and similar
tests at Mach numbers of 1.93 and 1.62 of a service
pitot-static tube to determine body static pressures
and indicated Mach numbers are presented and dis-
cussed. The radial pressure distribution on the cy-
lindrical bodies is compared with that calculated by
an approximate theory.


NACA-Langley 7-23-52 -4000




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