Research abstracts

CURRENT NACA REPORTS

NACA TN 2955

ESTIMATION OF FORCES AND MOMENTS DUE TO
ROLLING FOR SEVERAL SLENDER-TAIL CON-
FIGURATIONS AT SUPERSONIC SPEEDS. Percy J.
Bobbitt and Frank S. Malvestuto, Jr. July 1953.
71p. diagrs. (NACA TN 2955)

A method, based on conformal-transformation
techniques, for solving two-dimensional boundary-
value problems has been used to evaluate the velocity
potentials, span loadings, pressure distributions, and
associated stability derivatives for several slender-
tail arrangements performing a steady rolling mo-
tion. Illustrative variations of the rolling stability
derivatives for several series of tail shapes, as well
as sample span loadings and pressure distributions,
are included.

NACA TN 2967

AN ANALYSIS OF THE POWER-OFF LANDING
MANEUVER IN TERMS OF THE CAPABILITIES OF
THE PILOT AND THE AERODYNAMIC CHARAC-
TERISTICS OF THE AIRPLANE. Albert E. von
Doenhoff and George W. Jones, Jr. August 1953.
42p. diagrs. (NACA TN 2967)

A theoretical analysis of the power-off landing
maneuver is presented In which an attempt is made
to consider both the aerodynamic characteristics of
the airplane and the human capabilities of the pilot.
Certain factors are derived which allow for the var-
intions in pilor judgment in a manner which is di-
rectly connected to pilot capabilities; however, fu-
ture research Is needed for the actual numerical
evaluation of the factors. An analysis including
these factors is set up by means of which the mimi-
mum length of landing field required for landing an
airplane safely could be found.

NACA TN 2976

A STUDY OF THE STABILITY OF THE INCOM-
PRESSIBLE LAMINAR BOUNDARY LAYER ON IN-
FINITE WEDGES. Neal Telervin. August 1953.
41p. diagrs., 4 tabs. (NACA TN 2976)

Hartree's numerical solutions of the boundary-
layer equations for the flow over infinite wedges are
used to confirm a result prenlously obtained by the
use of Schlichting's approximate method for the cal-
culation of the laminar boundary layer; namely, that
in a region of falling pressure a thick velocity pro-
file can be more stable than a thin profile although
the velocity at the edge of the boundary layer and the


pressure grad~ent are the same for both profiles.
The Investlgation also leads to the inference that the
calculated effects of a change in boundary-layer
thickness on the stabusity and on thees
Reynolds number should beeseai
by replacing the Schlichting as qi ameter fiahqqy
of velocity profiles by the Har d ~single-parameter -
family of velocity profiles.
NACATNBBO ( DEC y 355

THE AERODYNAMIC CHARA E IC O A
ASPECT-RATIO-20 WING HA N
FOL SECTIONS AND EMPLOY BOUNDARY-.*
LAYER CONTROL BY SUCTION. B~me C ocke,
Jr., Marvin P. Fink and Stanley M. Gottlieb.
August 1953. 63p. diagrs., photos., 2 tabs.
(NACA TN 2980)

An investigation has been conducted to study the
aerodynamic characterishecs of an aspect-ratio-20
wing employing thiek airfoil sections and boundary-
layer control by suction. Data from models tested
In the Langley full-scale tunnel and the Langley low-
turbulence pressure.tunnel are included in this re-
port The results indicate the effects of varylng
suction flow rate, suction-slot configuration, wing
surface condition, fap deflection, and Mach number.

NACA TN 2982

SUPERSONIC FLOW PAST OSCILLATING AIRFOLS
INCLUDING NONLINEAR THICKNESS EF FECTS.
Milton D. Van Dyke. July 1953. 41p. diagrs.
(NACA TN 2982)

A solution to second order In thickness as derived
for harmonically ascillating two-dimensional airfoils
in supersonic flow. For slow osnilations of an
arbitrary profile, the result rs found as a series in-
cludang the third power of frequency For arbitrary
frequencies, the method of solution for any specific
profile Is indi~cated, and the exp~licit solution derived
for a single wedge. Nonianear thickness effects are
found generally to reduce the torsional damping, and
so enlarge the range of Mach numbers within which
torsional Instabdlity Is possible.

NACA TN 2985

A FLIGHT INVESTIGATION OF THE EF FECT OF
STEADY ROLLING ON THE NATURAL FREQUEN-
CIES OF A BODY-TAIL COMBINATION. Norman
R. Bergrun and Paul A. Nickel. August 1953 27p.
diagrs., photo 2 tabs. (NACA TN 2985)

Flight data have been obtained with a freely falling
body-tail combination to show the effects of steady


* AVAILABLE ON LOAN ONLY.
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 17P4 F ST, NW. WASHINGTON 9s. D C, CITING CODE NUMBER ABOVE EACH TITLE;
THE REPORT TITLE AND AUTHOR.


National Advisory Committee For Aeronautics


Research Abstracts


NO.47


AUGUST 19, 1953





RESEARCH; ABSTRACTS NO. 47


NAA TN 2992

APPLICATION OF A CHARACTERISTIC BLADE-
TO-BLAD)E SOLUTION TO FLOW IN A SUPER-
SONIC ROTOR WITH VARYING STREAM-FILAMENT
THIICKNESS. Eleanor L.Costlow. Auguatl 95.
36p. diagrs., 5 tabs. (NACA TN 2992)

An analysis of the circumferential blade-to-blade
flow properties in a supersonic impeller has been
made using the method of characteristics on an arbi-
trary stream surface of revolution. The method
takes into account variable stream-filament thick-
ness and curvatrire along the flow surface. The re-
sults of an example considering stream-filament-
thickness variation alone indicate an appreciable dif -
ference between the flow properties calculated with
the plane-flow characteristic equations and those de-
termined with the characteristics method demon-
strated here. The effect of a stream-filament-
thickness reduction from blade-passage inlet to exit
was to reduce the relative velocity and the absolute
value of the flow angle in the blade passage.

NACA TN 2993

CALIBRATION OF ST'RAIN-GAGE INSTALLATIONS
IN AIRCRA FT ST RUCT URES FOR THE MEASURE -
MENT OF FLIGHTI LOADS. T. R. Skopinslu,
William S. Aiken, Jr. and Wilber B. Huston.
August 1953. 70p. diagrs., 10 tabs. NACA
TN 2993. Formerly RM L52G31)

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

NACA RM E53E28

EXPERIMENTAL, INVESTIGATION OF THE EF-
FECTS OF SUPPORT INTERFERENCE ON THE
PRESSURE DISTRIBUTION OF A BODIY OF REVO
LUTION AT A MACH NUMBER QF 3. 12 AND) REY-
NOLDS NUMBERS FROM 2 x 10" to 14 x 106.
L. Eugene Baughman and John R. Jack. August
1953. 18p. diagrs., photo., tab. (NACA
RM E53E28)

An experimental investigation wasaeperformed to de-
termine the effect on base and forebody pressures of
using a sting Imodified wEith varying length splitter
plates and fins instead of a conventional sting to sup-
port a cone-cylinder body of revolution. The in-
vestigation was conducted at a Mach number of 3. 12
for a Reynolds number range of 2 x 106 to 14 x 106
and for an angle of attack range of Oo to So. For
Reynolds numbers of 8 x 106 and 14 x 106 there was
a negligible effect of the splitter plate modification
on the base pressure, and at a Reynolds number of
2 x 106 tere was asmall effect. Positioning the
leading edge of the splitter plate at or ahead of the
base made no appreciable change in the influence of
the modifications pn base pressure at a Reynolds
number of 14 x 10 With the fin-type modification
there was a small increase in base pressure.


rolling on the response in the pitching and yawng
motions. Observed effects are compared with those
predicted by theory, and reasonably good agreement
is obtained. Charts are also presented to provide
a rapid means for estimating the effective values of
the saii~yety deivatve Cm, and C"B under
steady rolling conditions.

NACA T`N 2988

TEMPERATURES, THERML STRESS, AND SHOCK
IN HEAT-GENERATING PLATES OF CONSTANT
CONDUCTIVITY AND OF CONDUCTIVITY THAT
VARIES LINEARY WITH TEMPERATURE. S. V.
Manson. July 1953. 62p. diagrs. (NACA
TN 2988)

Workig formulas are presented for the steady-state
temperatures and thermal stress in heat-generating
infinite plates of constant conductivity, and of con-
ductivity that decreases linearly with temperature
as the temperature increases, for the case in which
the heat is generated uniformly throughout the plate
thickness and both faces of the plate are equally
cooled. A criterion is indicated for determining the
surface cooling conditions under which the thermal
shocks at the surface and midiplane will be smaller
than, equal to, or greater than the steady-state
thermal stresses at those planes. The dimension-
less parameters governing the transient tempera-
tures and thermal stresses in materials of linearly
varying conductivity are derived by a similarity
study of the conduction equation and boundary con-
ditions of the transient state. A numerical tech-
nique for solving the transient-state equations is
indicated in detail.

N~ACA TN 2989

COMPARISON OF SECONDARY FLOWS AND
BOUNDARY-LAYER ACCUMULATIONS IN SEV-
ERAL TUJRBINE NOZZLES. Milton G. Kofakey,
Hubert W. Allen and Howard Z. Herzig. Augus'
1953. 58p. diagrs., photos., 3 tabs. (NACA
TN 2989)

An investigation was made of losses and secondary
flows in three different turbine nozzle configura-
tions in annular cascades. Appreciable outer
shroud loss cores (passage vortices) were found to
exist at the discharge of blades which had thickened
suction surface boundary layers near the outer
shroud. Blade designs having thinner boundary
layers did not show such outer shroud lose cores
but indicated greater inward radial flow of low mo-
mentum air in the blade wake which resulted in
greater contribution to inner shroud loss regions.
The blade wake was a combination of profile loss and
low momnentumn air from the outer shroud, and the
magnitude of the wake loss is to this extent an indi-
cation of the presence or absence of radial flow. At
a higher Mach number, shock-boundary-layer thick-
ening on the blade suction surfaces provided an
additional radial flow path for low mnomentumn air
which resulted in large inner shroud loss regions
accompanied by large deviations from design values
of discharge angle,






NACA
RESEARCH ABSTRACTS NO.47

NACA RM E53 FO2

PRELIMINARY RESULTS OF HEAT TRANSFER
FROM A STATIONARY AND) ROTATING ELLIPSOI-
DAL SPINNER. U. von Glahn. August 1953. 35P-
diagrs., photo., 2 tabs. (NACA RM E53FO2)

The convective heat -transfer coefficients were de-
termined for an ellipsoidal spinner of 30-inch maxi-
mum diameter for both stationary and rotating oper-
ation. The range of conditions studied included air-
speeds up to 275 miles per hour, rotational speeds
up to 1200 rpm, and angles of attack of zero and 4o-
The results rndicate that a higher heat tranfer oc-
curred with rotation of the spinner. Transition from
laminar to turbulent flow occurred over a large
range of Reynolds numbers primarily because of sur-
face roughness of the spinner.


BIRITISH REPORTS


N-23901

Royal Aircraft Establishment (Gt. Brit.)
FUNCTION GENERATORS BASED ON LINEAR IN-
TERPOLATION, WITH APPLICATIONS TO ANA-
LOGUE COMPUTING. E. G. C. Burt and O. Hi.
Lange. April 1953. 2p. diagrs. (RAE Tech. Note
GWV 244)

The use of function generators in electronic analogue
comrputing and simulation greatly extends the range
at problems which can be solved by these methods.
This paper presents a technique in which diode units
are used to approximate to the functions by linear in-
terpolation. It is shown that the method can be ex-
tended to deal wiith a wide class of functions, in-
cluding multvaiate functions. Analogue multiplica-
tion and division are discussed as particular cases
of funtion generators, and formulas for the general
function are developed. The results are presented
of an experimental generator for sin x in the range
- p Ex p, in which the error is about 1 percent of
the umamu output*

N-23902*

Royal Aircraft Establishment (Gt. Brit. )
MOVING; PARTIAL RANGE SMOOTHING WVITH
CENTRAL-DIFFERENCES USING A NATIONAL
ACCOUNTING MACHIN~E. J. D. Downes. April
1953. 33p. diagrs., 12 tabs. (RAE Tech.Note
GW 250)

Based on a suggestion by Dr. Th. W. Schmidt of
G. W. Trials Division, a method of partial range
smoohing of experimental data by central-
differences which greatly reduces the normal amount
of labor and permits simple checking, is described.
The method is particularly suited for use where a
National Accounting M~achine as well as a desk cal-
culator is available. The observed quantities are
converted to linear combinations of their difrences
and used in an adaptation of the 'least-squares"
method of curve-fitting. Additional, relevant appli-
cations of the National Machine are included in
nuerical exape.


N-23929e

Aeronautical Research Council (Gt. Brit. )
THE APPLICATION OF THE: POLY'GON METHO
TO THE CALCULATION OF THE COMPRESSIBLE:
SUBSONIC FLOW ROUND TWO-DIMENSIONAL
PROFILES. L.C. Woods. 1953. 31p. diagrs.,
7 tabs. (ARC CP 115)

This paper describes the method now used by the
author of applying the polygon method to the calcula-
tion of the compressible subsonic flow round two-
dimensional airfoils. Tables have been constructed
which can be used for all airfoil shapes, putting the
polygon method on the same footing numerically with
Goldstein's aApproximation IDI for incompressible
flow. A method of calculating lift and moment coef-
, ficients, and their rates of change with incidence is
also given in the paper.

N-23930*

Aeronautical Research Council (Gt. Brit. )
WAKE SURVEY AND STRAIN-GAUGE MEASURIE-
MENTS ON AN INCLINED PROPELLER IN THE
R. A. E. 24-FI'. TUNNEL. PART I. WAKE SUR-
VEY. J. G. Russell. 1953. 61p. diagrs., photos.,
4 tabs. (ARCCP 117)

This report describes tests carried out in the RAE
24-foot wind tunnel with a 16 foot diameter, 4 bladed
propeller, during October and November 1949.
Wake survey and blade strain-gage measurements
were made at tunnel speeds of 100 and 170 fps
with the propeller axis inclined at angles of Oo, 5o,
100, and 150 to the air flow. The blade angles and
propeller rotational speeds were also varied within
the limits imposed by the 1, 500 hp electric motor,
The lift grading curves at the points of maximum. and
minimum loading, derived from total head meas-
urements made in the slipstream by means of a pilot
comb, have been compared with estimated values,
and estimated power absorption figures compared
with measured values.

N-23931*

Aeronautical Research Council (Gt. Brit. )
A COMPANION OF CALCULATED AND MEAS-
URED BASE, PRESSURES OF CYLINDRICALLY
BASED PROJECTILES. W. F. Cope. Appendix:
CALCULATION OF REYNOLDS NUMBER EFFECT
ON PROJECTILES AT SUPERSONIC SPEEDS. W.
F. Cope. 1953. 12p. diagrs., photo. (ARC CP
118)

In 1946 at the VI International Congress for Applied
Mechanics a paper wras read on a method of calcula-
ting the base pressure of a cylindrically based pro-
jectile. At that time there were few written base
pressure determinations and therefore it was not
possible to check completely the theory put forward.
Since that time measurements have been made. In
this report these measurements are compared with
results of calculations according to the theory. The
original report, "Calculation of Reynolds Number
Effect on Projectiles at Supersonic Speeds, is in-
cluded as an appendi.




NACA
RESEARCH ABSTRACTS NO.47


This is an attempt to extend the usual blockage for-
mulas to high subsonic speeds. It la based on an
assumption which perhaps can only be fully justified
by results. The conclusion reached seems edIfl-
eiently reasonable to suggest that the method might
be worth taking further either analytically or ex-
perimentally.


N-23936*



ARECTU U VRE Sa he Mdathe~m ties
Division. 1953. 18p. (ARC R &M2461. Former-
ly ARC 10, 754; Perf. 384; & C '2138)

Critical tables are given from which the complete
downwash due to a unit rectangular vortex can be
read to three places of decimals, on selected lines
coplanar with the vortex. These lines are at speci-
fled integral multiples of the semlwidth of the vortex
from the center line of the vortex, and are of imr-
portance in the solution of wing loading problems by
vortex lattice theory.

N-23937*

Aeronautical Research Council (Gt. Brit.)
THE EFFECT OF CURVATURE OF SURFACE ANDI
THICKNESS OF TRAILING EDGE ON AILERO
HINGE MOMENTS. Part I. MODIFICATION TO
AILERON ON UPPER SURFACE. A. S. Batson,
C. H. Burge and J. R. Greening. Par II. MODIFI-
CATION TO AILRON ON LOWER SURFACE A.
S. Batson, C. H. Burge and W. C. Skelton. Part
II. MODIFICATION TO AILERON ON BOT SUR-
FACES. A. S. Batson, C. H. Burge and W. C.
S Ilton. 15350 48p~od Ar.,ht. 018tabs.127

ARC 5490; & C 1200; ARC 6188; 86 C 1426)

This investigation was made to provide additional
data on the effect on hinge moment of rounding the
upper surface and of thickening the trailing edge of
an aileron. Hinge moments were measured on an
aileron (24.6 percent balance), and on ain aileron
(27-1/2 percent balance) with various modifications
to the upper surface, lower surface, and both sur-
faces. Strips (approximately 1 percent of aileron
chord) were also fitted for a few cases near the
trailing edges.


N-23938*

Aeronautical Research Council (Gt. Brit.)
THE THEORETICAL ESTIMATION OF POWER
REQUIRMENTS FOR SLOT-SUCTION AEROFOILB,
WITH NUMERICAL RESULTS FOR TWO THICK
GRIFFITH TYPE SECTIONS. J. H. Preston,

(AR g M 25d7A{ orel 11@ 01 2.8027 gra.
276; FM 1052; ARC 11,610; Perf. 466; FM 1263)

This report describes a method for assessing the
performance of slot-suction airfoils in terms of an
effective drag coefficient, which takes into account
the power requirements of the suction pump neglect-
ing slot entry and duct losses. When the suction-alot
is located at a velocity discontinuity the suction flow


N-239324

Aeronautical Research Council (Gt. Brit.)
EFFECT OF MEAN STRESS ON THE FATIGUE
STRENGTH OF D. T.D. 364 ROUND BARS WITH AND
WITHOUT TRANSVERSE HOLES. G. M. Norris.
1953. 15p. photos., diagrs., 3 tabs. (ARC CP
120)

Endurance curves for zero mean stress and for ten-
sile mean stresses of 5 tons/in2 and 10 tons/in2
wr eermined tor round a eimennc em hined from

dettermined also for similr specimens with a trans-
verse hole. Variations in surface roughness caused
some scatter in results and an attempt is made to
correlate degree of surface finish with fatigue
strength.

N-23933+

Aeronautical Research Council (Gt. Brit. )
THE LOW SPEED PERFORMANCE OF A HELI-
COPTER. A. L.Oliver. 1953. 12p. diagrs.
(ARC CP 122)

The analysis and estimation of helicopter perform-
ance is dependent upon the accurate assessment of
rotor induced velocity. An empirical curve relating
the flow through the rotor to the flight speed is used
for vertical flight and the momentum theory is suf -
ficiently accurate for a tip speed ratio greater than
about 0. 1, but no simple method has been generally
available for the intermediary speed range. Sets of
empirical curves covering this speed range and
based on the analysis of low-speed flight perform-
ance are given in this report. The charts give
values of the rotor induced velocity varying smooth_
ly from the vertical flight state to the forward flight
region in which the momentum theory becomes ac-
curate. The charts are presented in forms suitable
for determining steady flight performance and also
for estimating rotor thrust during accelerated motion
in, for exaple, take-off flight. .


N-23934*

Aeronautical Research Council (Gt. Brit. )
THE USE OF INFLUENCE FACTORS IN PROBLEMS
OF FLUID FLOW. K. H.V. Britten. 1952. 13p.
diagrs., 2 tabs. (ARC R& M 2441. Formerly
ARC 10, 500; FM 1094)

A vigorous mathematical analysis has been applied to
some empirical reanlts, obtained by Thom, on the
use of the "squares method" in problems of fluid
flow. The problem considered is that of determining
the incompressible and compressible flow past an
arbitrarily shaped body placed in a wind tunnel.

N-23935*

Aeronautical Research Council (Gt. Brit. )
TUNNEL WALL ]EFFECT FROM MASS FLOW CON-
SIDERAIONS. A. Thom. 1953. 10p. diagrs., 2
tabs. (ARC R &M2442. Formerly AC11, 004;
FIM 11731 TP 211)




NACA
RESEARCH ABSTRACTS NO.4~7

required to prevent separation can be calculated,
using the elementary theory suggested by Sir
Geoffrey Taylor. The method is applied to two
Griffith type airfoils and the drags are compred with
those of normal thin airfoils 20 percent thick-

N-2393g"

Aeronautical Research Council (Gt. Brit.)
ELECTRONICS APPLIED TO THE MEASURMENT
OF PHYSICAL QUAN~T~ITIES G. E. Bennett, G. R.
Richards and E. C.Voss. 1952. 1214p. diagrs.,
photos. (ARC R M 2627. Formerly RAE
Instn. 1)

The report describes the application of electro-
mechanical and elect ronic principles to the design of
instruments for the measurement of physical quan-
tities such as movement, strain, pressure, accel-
eration, and vibratory motion, writh particular refer-
ence to the special requirements of aeronautical en-
gineering. The dynamic characteristics of pick-ups
are considered, and subdivided on an electrical
basis into electromagnetic, capacitance and resist-
ance types, a detailed description of each type being
given. An account is given of the circuits used for
the conversion of the electrical variation produced in
each type of pack-up into a corresponding voltage or
current, particular mention being made of bridge
circuits and resonance circuit methods. The spec-
ial requirements of amplifiers, and the best basic
circuits for satisfying them, are considered and il-
lustrated by detailed reference to a number of par-
ticular amplifier designs; in particular, direct-
coupled and carrier amplifiers are considered. The
requirements of recording equipment and the vari-
ous recording methods are discussed, and a de-
talled account given of photographic recording and
various oscillograph cameras, their optical arrange-
ments, components and timing devices.

N-23940

Aeronautical Research Council (Gt. Brit.)
TOWING TANK TESTS ON A LARGE SIX-ENGINE
FrLYING BOAT SEAPLAE, TO SPECIFICATION
10/46 PRINCESS. PART I. GENERAL PORPOISING
STABILITY, TRIM AND SPRAY CLEARANCE*
A. Gi. Smith, G. L. Fletcher, T. B. Owen and D). ~F.
Wright. 1953. 29p. diagrs., photos., 6 tabs.
( RR &M. 2641; 11, 462. Formerly RAE Aero


This report gives the results of the first series of
towing tank tests made at the Royal Aircraft Estab-
lishment Towing Tank (up to May 1947) on a powered
dynaic model of a six-engine transport flying boat,,
later naed the Princess class, and designed to
specification 10/46, on the basis of whch full-scale
hull construction was started; latter tests have been
made to further improve the hull step and afterbody
and test the effect of modifications to the aerodynam-
ic superstructure and power units.

N-26032*

National Gas Turbine Establishment (Gt. Brit.)
THE THERMIODYNAMYICS OF FRICTIONAL RE-
SISTED ADIABAIC FLOW OF GASES THROUGH
DUCTS O]F CONST~ANT AND VARYING CROSS SEC-
TION. W. R. Thomson. September 1952. 45p.
diagram. (NGTE R.110)


The report presents an analytical study dealing with
the adiabatic flow of gases with frictional losses
through ducts of constant and varying cross section.
The thermodynamic treatment is along lines pub-
lished by other workers such as Bailey and Fabri and
is essentially one-dimensional in character insofar
that frictional effects are assumed to be uniformly
distributed over the total cross-sectional area of
flow. With this simplifying assumption, relation-
ships are deduced connecting the pressure, tempera-
ture, velocity and flow area of the gas at any one
plane wcith those at any other plane in a duct. The
main relationships are unusable for quantitative es-
timation except through graphs and the main value of
the report lies in the presentation of these graphs,
the use of which should facilitate the solution of duct
flow problems.


N-26067*

Aeronautical Research Council (Gt. Brit. )
ON BOUNDARY LAYERS AND UPSTREAM IN-
FLUENCE. I. A COMPARISON BETWEEN SUB-
SONIC AND SUPERSONIC FLOIWS. M. J. Lighthill.
October 23, 1952. 17p. diagrs., 3 tabs. (ARC
15, 297; FM 1805)

It is pointed out that there are two separate mechan-
isms for upstream influence through the boundary
layer in supersonic flow, and that one of these (that
involving separation) operates also in subsonic flow.
A quantitative theory of subsonic flow up a step is
given to illustrate this. The main differences be-
tween the subsonic and supersonic flows are as
follows. (1) The boundaries of dead air regions are
nearly straight in supersonic flow but are usually
highly curved in subsonic flow. (2) Separation
(whether of the laminar or turbulent layer) occurs at
a much lower pressure coefficient in supersonic flow;
this is only slightly due to the fact that the fluid near-
est the wall is then lighter and so more easily brought
to rest; it is due much more to the relative sudden-
ness of the pressure rise ahead of the dead air region.
(3) However, for a given pressure coefficient in the
dead air region, the distance of upstream influence
is somewhat greater in the subsonic flow, except at
the higher pressures.


N-26085*

National Gas Turbine Establishment (Gt. Brit. )
A METHOD FOR MEASURING THE REACTION
FORCE OF A HIGH PRESSURE GAS STREAML. P.
J. Fletcher. March 1953. 9p. diagrs. (NGTE
Memo. M. 184)

A method has been developed for measuring the re-
action force produced by a high pressure gas stream.
Briefly the method is to utilize the jet reaction to
produce a bending moment in the gas supply pipe by
arranging for the jet to discharge in a direction at
right angles to that of the entering air. The strain
due to bendig is indicated by resistance strain gages
and the reaction force deduced from direct calibra-
tion. The method thus obviates the need for flexible
joints and hence eliminates the problem of ensuring
flexibility and sensitivity at high pressure. A de-
tailed description of the method of installing the
gages is given as well as the precautions necessary







NACA
RESEARCH ABSTRACTS NO.47


to separate the required bending strain from that due
to pressure forces. With the particular apparatus
described a thrust of up to 150 lb produced by a gas
stream at 10 atm pressure can be measured to writh-
in +0. 21b.


NACA-Langley 8-19-53- 4000

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