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:
AA00009235:00090

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Research Abstracts 0 /
SEPTEMBER 16, 1952


CURRENT NACA REPORTS


NACA Rept. 1028

EFFECT OF ASPECT RATIO ON THE AIR FORCES
AND MOMENTS OF HARMONICALLY OSCILLATING
THIN RECTANGULAR WINGS IN SUPERSONIC PO-
TENTIAL FLOW. Charles E. Watkins. 1951. 17p.
diagrs. (NACA Rept. 1028. Formerly NACA
TN 2064)

The linearized velocity potential for a thin, flat, rec-
tangular wing undergoing sinusoidal torsional oscil-
lations simultaneously with sinusoidal vertical trans-
lations in a supersonic stream is derived in the form
of an infinite series in terms of a frequency param-
eter. Simple closed expressions that include the re-
duced frequency to the third power are given for the
velocity potential, components of total force and mo-
ment coefficients, and components of section force
and moment coefficients. It is found that the com-
ponents of force and moment coefficients for small-
aspect-ratio wings may deviate considerably from
those for an infinite-aspect-ratio wing. The method
of solution can be utilized for other plan forms, that
is plan forms for which steady-state solutions are
known.


NACA Rept. 1053

INVESTIGATION OF TURBULENT FLOW IN A TWO-
DIMENSIONAL CHANNEL. John Laufer, California
Institute of Technology. 1951. 20p. diagrs., photos.
(NACA Rept. 1053. Formerly TN 2123)

A detailed exploration of the turbulent flow character-
istics in a two-dimensional channel is presented.
The measurements were made at three Reynolds
numbers, 12,300, 30,800, and 61,600, based on the
half width of the channel and the maximrhum mean
velocity. A channel of 5-inch width and 12:1 aspect
ratio was used for the investigation.



NACA TN 2748

ON TRANSONIC FLOW PAST A WAVE-SHAPED
WALL. Carl Kaplan. August 1952. 43p. diagrs.
(NACA TN 2748)

The simplified nonlinear differential equation for
transonic flow past a wavy wall is solved by the meth-


od of integration a for
'se'el lger4i'.ip teurefo
the solution of tesilting recurrence rr-inhlas is
shown and illust i% oy a number of eAx.'wes A
numerical test of -ln".'-'etence is applied ti a key
poAer series in k. Bmnlirn '1e S -i larXyr paran, -
eter. and leads to the cluf tnibn t- 'iooth sym-
metrical potential flow pa -- 'y wall is no longer
possible when the critical value of the stream Mach
number is exceeded.


NACA TN 2749
1
ANALYSIS OF FLOW IN A SUBSONIC MIXED-FLOW
IMPELLER. Chung-Hua Wu, Curtis A. Brown and
Eleanor L. Costilow. August 1952. 38p. diagrs.
(NACA TN 2749)

A method recently developed for determining the
steady flow of a nonviscous compressible fluid along
a relative stream surface extending from hub to cas-
ing between two adjacent blades in a turbomachine is
applied to investigate the through flow of air in an
experimental mixed-flow impeller of high solidity.
The shape of the stream surface is taken to be the
same as that of the mean camber surface of the blade
which consists of all radial elements. The principal
equation governing the through flow is solved by the
relaxation method with the use of fourth-degree
differentiation formulas for unequally spaced grid
points caused by the varying hub and casing wall radii.
A detailed analysis is made of both incompressible
and compressible flow through the impeller, and
contour plots of the stream function, velocity com-
ponents, total enthalpy, static pressure, and Mach
number are presented and discussed.


NACA TN 2762

AERODYNAMIC CHARACTERISTICS OF THREE
DEEP-STEP PLANING-TAIL FLYING-BOAT HULLS
AND A TRANSVERSE-STEP HULL WITH EXTENDED
AFTERBODY. John M. Riebe and Rodger L.
Naeseth. August 1952. diagrs., photos., 5 tabs.
(NACA TN 2762. Formerly RM L8127)

An investigation was made to determine the aerody-
namic characteristics of three deep-step (92 percent
of beam) planing-tail flying-boat hulls differing only
in the amount of step fairing and of a transverse-step
hull with extended afterbody. Minimum drag coeffi-
cients, which include the interference effects of the
support wing, were 0. 0066 for the transverse-step
hull about the same as for a conventional hull and


*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. CLASSIFIED DOCUMENT
I hs doument contains classified itnforrmation afecting the Nattonal Defense of the United StatI w, trn the Inlb, ato n :.tn cqa tt to .y r' -I t' r o period rn th a m l ry arlo ova. s; a ;'*'. )a
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t an Unauthortied person Is prohibited by law. I Uhite s.Ictat zens o' knof s loyalty .m sreUh h o f ntcesl ul tf. n..r et3 !te .


N0.O29


National Advisory Committee for Aeronautics








2


0. 0057 or 14 percent less for the hull with a deep un-
faired step. The hulls with step fairing had up to 44
percent less drag. Lii,,iluIhi.il and lateral stability
was about the same as for a conventional hull.


NACA TN 2763

GUST-RESPONSE ANALYSIS OF AN AIRPLANE IN-
CLUDING 'A ING BEND[NLu FLEXIBILITY. John C.
Houbolt and Eldon E. Kordes. August 1952. 48p.
di.irs., 3 tabs. (NACA-TN 2763)

An i.il.'is is made of the gust response (including
bending mnioment) of an airplane ha% ii, the degrees of
freedom of vertical motion and wing I.,lnd.li.i' ii \iili-
ity and basic parameters are established. A con-
venient numerical solution of the response equations,
well-suited for trend studies, is developed and used
in an example. A method is indicated for determin-
ing the gust causing a known response and a pro-
cedure is given for determining the response of an
airplane directly from the known response of another
airplane.


NACA TN 2766

SOME EFFECTS OF AMPLITUDE AND I ILQUENCY
ON THE AERODYNAMIC DAMPING OF A MODEL
OSCILLATING CONTINUOUSLY IN YAW. Lewis R.
Fisher and Walter D. Wolhart. September 1952.
24p. diagrs., photo. (NACA TN 2766)


A fuselage-vertical-tail combination was oscillated
in yaw through a range of amplitudes from 1 2 to
4 and a low range of the reduced-frequency param-
eter. The phase angles between the tail force and
the displacement were measured and converted to
values of the damping in yaw which are compared
with the damping predicted by the unsteady-lift theo-
ries and with the experimental steady-state daniping
value. These tests were conducted at a Reynolds
number of 442, 000.


NACA TN 2769

EXPERIMENTAL AND THEORETICAL DETERMI-
NATION OF THERMAL STRESSES IN A FLAT
PLATE. Richard R. Heldenfels and William M.
Roberts. August 1952. 35p. diagrs., photo.
(NACA TN 2769)

Thermal stresses induced in a flat, rectangular,
75S-T6 aluminum-alloy plate by nonuniform heating
are determined both experimentally and theoretically.
The characteristics of commercially available bond-
ed resistance wire strain gages are first investigated
to determine their suitability for measuring stresses
under simple conditions of stress and temperature.
The gages are then used to measure thermal stresses
in the flat plate in order to study their suitability
under more complicated conditions. The experi-
mental results are found to be in satisfactory agree-
ment (within t5 percent of maximum calculated
stress) with an approximate theoretical solution of
the problem.


NACA
RESEARCH ABSTRACTS NO.29


NACA TN 2770

STUDY OF THE PRESSURE RISE ACROSS SHOCK
WAVES REQUIRED TO SEPARATE LAMINAR AND
TLRRBLiLENT BOUNDARY LAYERS. ColemanduP.
Donaldson and Roy H. Lange. St-ptenet-r 1952. 20p.
diagrs., photos., tab. (NACATN 2770. Formerly
RM L52C21)

Results are presented of a dimensional study and an
experimental investigation of the pressure rise
across a shock wave which causes separation of the
boundary layer on a flat plate. The experimental
part of the investigation was conducted at a Mach
number of 3. 03 for a Reynolds number range of
2 x 106 to 19 x 106. The available experimental data
are compared with the predictions of the present
studi,., and the >iLiifi in, e of the results obtained is
discussed relative to certain practical design prob-
lems.


NACA TN 2771

THERMAL BUCKLING OF PLATES. Myron L.
Gossard, Paul Seide, and William M. Roberts.
August 1952. 39p. diagrs. (NACA TN 2771)

An approximate method, based on large-deflection
plate theory, for calculating the deflections of flat or
initially imperfect plates subject to thermal buckling
is outlined. The method is used to determine the de-
flections of a simply supported panel subjected to a
tentlike temperature distribution over the plate sur-
face. Experimental results for a particular panel
are in good agreement with the theoretical results
considered in the test.


NACA TN 2772

DRIVING STANDING WAVES BY HEAT ADDITION.
Perry L. Blackshear, Jr. August 1952. 46p.
diagrs., photos. (NACA TN 2772)

Types of burner instability are enumerated and the
role of standing waves in burners is discussed. The
status of the problem of flame-driven standing waves
is reviewed and a one-dimensional flow theory giving
the mechanism whereby a flame drives or damps a
standing wave is presented. In this theory, the re-
flection, transmission, and amplification of waves
passing through a flame region were determined from
the continuity and momentum equations. For the
model considered, waves were found to pass through
the flame front with their velocity amplitudes unal-
tered so long as the flame area remained unchanged.
A change in flame area acted as a source of waves
propagating simultaneously into the hot and cold gases
on either side of the flame zone. The one dimension-
al theory seems an adequate explanation of the exper-
imental observations.







NACA
RESEARCH ABSTRACTS NO.29


NACA TN 2774

A METHOD FOR FINDING A LEAST-SQUARES
POLYNOMIAL THAT PASSES THROUGH A SPECI-
FIED POINT WITH SPECIFIED DERIVATIVES.
Neal Tetervin. September 1952. lip. (NACA
TN 2774)

A method is presented for finding a mth-degree poly-
nomial that passes through a specified point with
(m j) specified derivatives (1 = j < m) and is a
least-squares polynomial for other points which are
spaced at unequal intervals of the independent vari-
able.


NACA TN 2775

EFFECT OF LINEAR SPANWISE VARIATIONS OF
TWIST AND CIRCULAR-ARC CAMBER ON LOW-
SPEED STATIC STABILITY, ROLLING, AND YAW-
ING CHARACTERISTICS OF A 45 SWEPTBACK
WING OF ASPECT RATIO 4 AND TAPER RATIO 0.6.
Byron M. Jaquet. August 1952. 27p. diagrs.,
2 tabs. (NACA TN 2775)

An investigation was made in the Langley stability
tunnel at low scale to determine low-speed effects of
linear spanwise variations of twist and circular-arc
camber on static-stability and rotary-stability (roll-
ing and yawing) derivatives of a 45 sweptback wing
of aspect ratio 4 and taper ratio 0.6.


NACA TN 2776

THE EFFECT OF A SIMULATED PROPELLER
SLIPSTREAM ON THE AERODYNAMIC CHARAC-
TERISTICS OF AN UNSWEPT WING PANEL WITH
AND WITHOUT NACELLES AT MACH NUMBERS
FROM 0. 30 TO 0. 86. Gareth H. Jordan and Richard
I. Cole. September 1952. 15p. diagrs., photo.
(NACA TN 2776)

Results of an investigation are presented to show the
effect of a simulated propeller slipstream on the lift,
drag, and pitching-moment characteristics of an un-
swept wing panel with and without nacelles. The
data were obtained in the Langley 24-inch high-speed
tunnel through a range of Mach numbers from ap-
proximately 0. 30 to 0. 86 for angles of attack of 0
and 3.



NACA TN 2777

THEORETICAL DISTRIBUTION OF SLIP ANGLES
IN AN AGGREGATE OF FACE-CENTERED CUBIC
CRYSTALS. John M. Hedgepeth. August 1952.
32p. diagrs. (NACA TN 2777)

An analysis of the relative frequency of occurrence
of any given slip-line angle in a plastically deformed
polycrystal composed of face-centered cubic crystals
is presented for the case of simple tension. The re-
sults are compared with those obtained for a poly-
crystal composed of crystals which have but a single


3


mode of slip and with experimental results. The
comparisons show that the differences between the
results obtained by the two theories become greater
as the stress is increased. The comparison with
experiment of the face-centered cubic theory is
somewhat better than that of the single-slip-mode
theory, but the errors are appreciable.


NACA TN 2778

STRESS AND STRAIN AT ONSET OF CRAZNLu OF
POLYMETHYL METHACRYLATE AT VARIOUS
TEMPERATURES. M. A. Sherman and B. M.
Axilrod, National Bureau of Standards. September
1952. 21p. diagrs., photos., 3 tabs. (NACA
TN 2778)

Stress and strain at the onset of crazing were deter-
mined for both general-purpose and heat-resistant
grades of commercial cast polymethyl methacrylate.
Tests were made at 230, 50, and 70 C on samples
0. 15 inch thick. Other properties measured were
tensile -ti enth. total elongation, and modulus of
elasticity. Results obtained indicate that a "critical-
strain theory" for the threshold of crazing is not ap-
plicable to polymethyl methacrylate. Strain at the
onset of crazing decreased with increase in tempera-
ture from 23 to 50 C whereas there was no consist-
ent trend from 50 to 70 C. Stress at the onset of
crazing was about 80 to 95 percent of the tensile
strength at all three temperatures.


NACA RM E52F19

VISUALIZATION OF SECONDARY-FLOW PHENOME-
NA IN BLADE ROW. H. Z. Herzig, A. G. Hansen
and G. R. Costello. August 1952. 20p. photos.,
diagrs. (NACA RM E52F19)


Flow-visualization methods were applied in a prelim-
inary study of the streamline pattern of secondary
flows in a blade row. The investigation demonstrated
flow of the inlet-wall boundary layer in a blade pas-
sage into the corner between the blade suction surface
and the wall to form a vortex. This vortex formation
occurs well within the passage and is not a trailing-
edge phenomenon. The magnitudes of the velocity
gradients and angle deflections make the significance
of quantitative measurements in this region question-
able.


NACA RM E52G17

EFFECT OF FREE METHYL RADICALS ON SLOW
OXIDATION OF PROPANE AND ETHANE. Glen E.
McDonald and Rose L. Schalla. August 1952. 21p.
diagrs., 4 tabs. (NACA RM E52G17)

A study of the effect of free methyl radicals on the
slow oxidation of both ethane and propane was made
by means of a photochemical decomposition reaction
in a static system. The introduction of methyl radi-
cals into a mixture of propane and oxygen or ethane
and oxygen substantially lowered the initiation tem-







4


perature of the combustion of the hydrocarbon.
Measurements were made of the amount of propane
oxidized in the presence of free methyl radicals at
temperatures ranging from 0 to 116 C and for two
different concentrations of propane. Oxidation of
ethane, which was induced by the free nietii.l radical,
was investigated for only one concentration from 0 to
183 C. From a plot of the log percentage of propane
oxidized against the reciprocal temperature (OK), a
zero activation energy, was found for the reaction be-
low 400 C. For ethane the activation eiieriV was
zero below 120 C. Above these two temperatures an
increase in the activation energy was observed. A
mechanism is proposed to account for the zero acti-
vation en-ertr",, but no attempt was made to interpret
the .iihc temperature reactions.


NACA RM E52G24

INFLUENCE OF EXTERNAL VARIABLES ON SMOK-
ING OF BENZENE FLAMES. Thomas P. Clark.
August 1952. 17p. diagrs. (NACA RM E52G24)

Premixed benzene-air flames were burned with vari-
ations in initial gas temperature, fuel-flow rate,
flame length, secondary-air-flow rate and burner-
tube diameter. Fuel-air ratios were measured both
at the initial appearance of incandescent carbon in the
flame and at the point where smoke issued from the
flame. Temperature alone affected the fuel-air ratio
at which carbon first formed in the flame. The
flame did not smoke under a fuel-air ratio of 164 per-
cent of stoichiometric, i-d ,lltss of the environment
around the flame. Conditions which enhanced the dif-
fusion of ,:,.,'t into the outer cone improved the
smoke-burning characteristics of the flame.


NACA TM 1343

THE EXCITATION OF UNSTABLE PERTURBATIONS
IN A LAMINAR FRICTION LAYER. (Die Anfachung
instabiler Storungen in einer laminaren Reibungs-
schicht). J. Pretsch. September 1952. 63p.
diagrs., 3 tabs. (NACA TM 1343. Trans. from
Aerodynamische Versuchsanstalt Gottingen E. V.,
Institute fur Forschungsflugbetrieb und Flugwesen;
Jahrbuch der deutschen Luftfahrtforschung, August
1942, p. I 54-71).

With the aid of the method of small oscillations which
was used successfully in the investigation of the
stability of laminar velocity distributions in the
presence of two-dimensional perturbations, the
excitation of the unstable perturbations for the Hartree
velocity distributions occurring in plane boundary-
layer flow for decreasing and increasing pressure is
calculated as a supplement to a former report. The
results of this investigation are to make a contribu-
tion toward calculation of the transition point on
cylindrical bodies.


NACA TM 1348

ON THE REPRESENTATION OF THE STABILITY
REGION IN OSCILLATION PROBLEMS WITH THE
AID OF THE HURWITZ DETERMINANTS. (Zur


NACA
RESEARCH ABSTRACTS N0.29


Darstellung des Stabilitatsgebietes bei Schwingungs-
aufgaben mit Hilfe der Hurwitz-Determinanten). E.
Sponder. August 1952. 12p. diagrs. (NACA
TM 1348. Trans from Schweizer Archiv fur ange-
wandte Wissenschaft und Technik, v. 16, no. 3, March
1950, p. 93-96).

This report concerns the use of the Hurwitz deter-
minants in defining boundaries of regions where os-
cillatory phenomena are to be stable or unstable. A
simplification is suggested as an aid in reducing the
computations usually required, although it is empha-
sized that point checks in the various regions defined
are required using the complete set of Hurwitz de-
terminants or some other complete stability deter-
mination.


NACA TM 1352

TRANSITION CAUSED BY THE LAMINAR FLOW
SEPARATION. (S6ryu-Hakuri ni tomnionau Sen'i ni
kansuru Kenkyu). T. Maekawa and S. Atsumi.
September 1952. 14p. diagrs., 2 tabs. (NACA
TM 1352. Trans from Society of Applied Mechanics
of Japan, Journal, v. 1, no. 6, November 1948,
p. 187-192)

An experimental investigation of the effects of the
geometry of body surface, Reynolds number, stream
turbulence, and a roughness element (wire) on the
reattachment of separated laminar boundary-layer
flow on a bent flat plate is presented and discussed.
The flow mechanisms determining reattachment of
the boundary layer are analyzed and discussed.



BRITISH REPORTS


N-16809*

Aeronautical Research Council (Gt. Brit.)
GROUND EFFECT ON DOWNWASH WITH SLIP-
STREAM. P. R. Owen and H. Hogg. 1952. 12p.
diagrs., 8 tabs. (ARC R M 2449; ARC 7590.
Formerly RAE Aero 1901)

It is shown in this report that given the downwash at
the tail plane with slipstream, then the decrement in
downwash at the same incidence and propeller thrust
coefficient due to ground interference can be esti-
mated by the formula of Piper and Davies (with cer-
tain slight modifications) with sufficient accuracy for
most practical purposes.



N-16810*

Aeronautical Research Council (Gt. Brit.)
TESTS ON A CONTROL WITH FIXED TRAILING-
EDGE ANGLE, AND SURFACES OF VARIABLE
CAMBER. A. S. Halliday, H. Deacon and W. C.
Skelton. 1952. 15p. diagrs., 3 tabs. (ARC
R & M 2495. Formerly ARC 10, 598; S &C 2116)







NACA
RESEARCH ABSTRACTS NO.29


The report describes experiments to determine the
effect of surface distortion on the hinge moments of a
control and to investigate the importance of a rigid
trailing edge in the prevention of large changes in
hinge-moment coefficient. The results show that by
designing the control so that the surface at the trail-
ing edge, extending over one-eighth of the chord, is
incapable of distortion, then the value of b2 remains
almost constant within the limits on any possible dis-
tortion forward of the fixed portion. It has also been
found that bI is more constant. Changes in center
line camber of the deformed portion of the control can
still produce an effective change in b2 when the cam-
ber varies with the control angle, although the amount
of this change may be roughly halved. Present prac-
tice of treating trailing-edge angle as an important
parameter in specifying control characteristics is
well justified.



N-16811'*

Aeronautical Research Council (Gt. Brit.)
LIMITATIONS OF USE OF BUSEMANN'S SECOND-
ORDER SUPERSONIC AEROFOIL THEORY. W. F.
Hilton. 1952. 9p. diagrs. (ARC R & M 2524.
Formerly ARC 9869; FM 971)

The author has found the Busemann theory very rapid
in use for the determination of CL, CD, and Cm,
and this report will enable the exact scope of its use
to be determined. It has been tacitly assumed in the
past that Busemann's second-order theory of airfoils
at supersonic speeds was subject to the same limita-
tions of wedge angle as the exact theory given by
Lighthill and others, namely, the wedge angle at
which the bow wave detaches. The range of angles
for which Busemann's theory gives a pressure coef-
ficient in error by less than 1 percent is shown to be
smaller than the angle range for the shock wave to be
attached. There is also a limit to the application of
Busemann's method to angles of expansion as well as
to angles of compression, unlike the exact theory,
which can be extended to expansive angles of the
order of one right angle without breaking down, in
fact far beyond the useful range. The limits of angle
given for the use of Busemann's theory are conserv-
ative, since they give the pressures to 1 percent,
and the force coefficients will be more accurately de-
termined since the errors tend to cancel out when
integrating pressures to obtain forces.



N-16812*

Aeronautical Research Council (Gt. Brit.)
ABSTRACTS OF PAPERS PUBLISHED EXTERN-
ALLY. 1952. 17p. (ARC R & M 2565)

Presents an index and related abstracts of papers
discussed by the Aeronautical Research Council and
recommended by them for external publication.


5


N-16813"

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE EFFECT OF SIZE OF AIRCRAFT
UPON THE DIFFICULTIES INVOLVED IN LANDING
AN AIRCRAFT. D. Adamson. 1951. 13p. diagrs.,
tab. (ARC R & M 2567; ARC 10,903. Formerly
RAE Aero 2202)

The effect of aircraft size upon the response of an
aircraft during those maneuvers which are commonly
employed in landing has been examined, and in this
way an assessment has been made of the way in vhich
the difficulties experienced by the pilot will go up as
aircraft size increases. On the basis of the work
summarized in this note, it is concluded that the
problems associated with landing (from the pilot's
point of view at any rate) are unlikely to be aggrava-
ted to such an extent, as the size of aircraft increases
up to the limiting size considered in this report (300-
ft span), as to make landing really difficult.


N-16814*

Aeronautical Research Council (Gt. Brit.)
ON THE MOMENTUM EQUATION IN LAMINAR
BOUNDARY-LAYER FLOW. A NEW METHOD OF
UNIPARAMETRIC CALCULATION. B. Thwaites.
1952. 9p. (ARC R & M 2587. Formerly ARC
11,155; FM 1193; Perf. 397)

The general method of Pohlhausen, discussed in de-
tail in reference 1, uses a uniparametric system of
velocity distributions of the form u/U = f(y/ 6) +
xg(y/5). Pohlhausen, by choosing simple forms for
the functions f and g, then uses the momentum
equation to find the distribution of 6 with x and
thence the distributions with x of the other bound-
ary layer characteristics. Several awkwardnesses
exist in his method, especially when it is applied to
problems dealing with a normal velocity at the bound-
ary. In this paper, a new method is described of
combining velocity distributions in the form
y/O = F(u/U) + AG(u/U), and it is shown that such a
combination avoids several difficulties. This method
of combination also allows a second parameter apart
from X, which might be found valuable in certain
problems. The method has been briefly described
before as part of an investigation into the effect of
continuous suction on laminar boundary-layer flow
under adverse pressure gradients. In that paper
(R & M 2514), a numerical example of its use was
given. In this paper, no example will be given be-
cause, as far as the author can see, the practical
use of the method is superseded by the generalized
method of reference 1; however, it possessed con-
siderable analytical interest.


N-16815*

Aeronautical Research Council (Gt. Brit.)
COMPARATIVE TESTS OF THICK AND THIN TURN-
ING VANES IN THE ROYAL AIRCRAFT ESTABLISH-
MENT 4 x 3-FT WIND TUNNEL. K. G. Winter.
1952. 5p. diagrs., tab. (ARC R & M 2589; ARC
10,976. Formerly RAE Aero 2217)








6


The tests were made by replacing the existing center
six thick vanes at the first corner of the 4- x 3-ft
wind tunnel by vanes of sheet metal. The thin vanes
reduced the corner loss, estimated from a wake
traverse behind one vane, without any deterioration
in outflow, and are therefore recommended for use in
future wind tunnels.


N-16816*

Aeronautical Research Council (Gt. Brit.)
TABLES OF MULTHOPP AND OTHER FUNCTIONS
FOR USE IN LIFTING-LINE AND LIFTING-PLANE
THEORY. V. M. Falkner. WITH APPENDIX:
CALCULATION OF THE CIRCULATION FROM THE
DOWNWASH. E. J. Watson. 1952. 52p. diagrs.,
8 tabs. (ARC R & M 2593. Formerly ARC 11, 234;
S & C 2187; Perf. 405)

The report gives the derivation and computed tables
of two classes of functions suitable for the solution
of problems of spanwise aerodynamic loading of
wings either by lifting-line or lifting-plane theory.
The functions are based on lifting-line theory, but,
by a consideration of the connection between lifting-
line and lifting-plane theory through the application
of Munk's stitit: r theorem to the calculation of in-
dueed (I rai. it is deduced that the functions must be
equally suitable for lifting-plane theory. The first
range of functions, called Multhopp or M functions, is
associated with discontinuities of induced downwash,
while the second, called P functions because of the
p.,1%. v- -ial representation of induced downwash, is
connected with discontinuities in rate of change of
induced downwash. Examples are given of the com-
bination of functions to produce given curves of in-
duced downwash, and evidence of the close relation
between the results for a continuous and stepped
downwash curve suggests that the functions tabulated
will be sufficient to cover almost any problem in
wing loading.



N-16817*

Aeronautical Research Council (Gt. Brit.)
CALCULATED AERODYNAMIC CHARACTERISTICS
OF TWO INFINITE WINGS WITH CONSTANT CHORD.
V. M. Falkner. 1952. lOp. diagrs., 8 tabs. (ARC
R & M 2594. Formerly ARC 10,628; S & C 2123;
Perf. 325)

The report gives solutions obtained by the vortex lat-
tice method for the aerodynamic loading of two infi-
nite wings of constant chord with sweepback of 45,
one with a V-joint at the center, the other rounded off
with arcs of radius four times the chord. The true
mathematical solution for these problems is exceed-
ingly difficult to find, and the accuracy has been veri-
fied by considering the convergence of solutions of
varying complexity. The V-wing shows a reduction
in circulation near the joint with accompanying back-
ward movement of the local center of pressure, while
the rounded wing has increased circulation without
appreciable variation of the center of pressure from
the 0. 25-chord position. The results will be used to


NACA
RESEARCH ABSTRACTS NO.29


modify loading functions used in vortex lattice theory
in order to improve solutions for wings of small as-
pect ratio, particularly when the leading or trailing
edges meet at an included angle which differs consid-
erably from 180.


N-1681b*

Aeronautical Research Council (Gt. Brit.)
APPLICATION OF THE LINEAR PERTURBATION
THEORY TO COMPRESSIBLE FLOW ABOUT BODIES
OF REVOLUTION. A. D. Young and S. Kirkby.
1952. 8p. diagrs. (ARC R & M 2624; ARC 11,033;
FM 1177. Formerly College of Aeronautics, Cran-
field Rept. 11)

The linearized theory is developed in some detail in
order to clarify the difference between two dimension-
al flow and flow past slender three-dimensional bod-
ies. In agreement with other authors, it is conclud-
ed that the perturbation velocity on a slender body in
compressible flow is 1/'62 times the perturbation ve-
locity in incompressible flow on a thinner body at re-
duced incidence obtained by reducing the lateral di-
mensions of the original body in the ratio 3:1. This
result is applied to a representative family of stream-
line bodies of revolution at zero incidence. It is
found that, without an undue loss of accuracy, the re-
sults of the claculations can be presented in a rela-
tively simple form in a diagram showing the variation
of velocity with Mach number for a range of values of
velocity on the surface of a streamline body in incom-
pressible flow. This variation is always less than
that predicted by the Glauert law but approaches it
with increase in the basic incompressible flow veloc-
ity, being very close to it for basic incompressible
velocity ratios, u/U0, of 1.10 and higher. It is
shown that the blockage factor for a body of revolu-
tion in a wind tunnel is increased in compressible
flow in the ratio 1/33 and not 1/i34 as quoted in
R &M 1909.


N-16819*

Aeronautical Research Council (Gt. Brit.)
MODEL TESTS IN THE 24-FT WIND TUNNEL TO
DETERMINE THE OPTIMUM ANGLE FOR RELEASE
OF A COCKPIT HOOD. R. Fail. 1952. 3p.
diagrs. (ARC R & M 2644; ARC 11, 494. Formerly
RAE Tech. Note Aero 1947)

For some time now, it has been recommended that
mechanical assistance be incorporated in jettisonable
cockpit hood designs. Some firms have preferred
designs in which the hood is constrained to rotate
through a definite angle before the final release. A
short series of tests has, therefore, been made in the
24-ft wind tunnel to determine the optimum angle for
release. This was found to be about 10, which is
considerably less than has been suggested in the past.

N-16820*

Aeronautical Research Council (Gt. Brit.)
WIND-TUNNEL TESTS ON A THICK SUCTION AERO-
FOIL WITH A SINGLE SLOT. M. B. Glauert, W. S.
Walker, W. G. Raymer and N. Gregory. 1952.







NACA
RESEARCH ABSTRACTS NO.29


14p. diagrs., tab. (ARC R &M 2646; ARC 11,797.
Formerly ARC 10, 854; Perf. 348; FM 1150)

This report describes the two-dimensional wind-
tunnel tests carried out in the National Physical Lab-
oratory 13- by 9-ft wind tunnel on a 31. 5-percent-
thick suction airfoil, GLAS II, which has a single slot
on the upper surface at 69 percent chord. Both suc-
tion and blowing were used to prevent separation.
Lift, drag, pitching moment, and the flow through the
slot were measured. Tests without suction were
made at Reynolds numbers of 0. 96 and 2. 8b million.
The results at the two Reynolds numbers were mark-
edly different, and at the higher speed widely varying
values of the drag coefficient were recorded in the
same conditions, there apparently being several pos-
sible regimes of flow. With suction, the pump power
available only enabled tests to be made at the lower
Reynolds number, and with the boundary layer on the
upper surface laminar to the slot. At low incidences,
suction quantities agreeing well with theoretical esti-
mates sufficed to maintain unseparated flow, but at
higher incidences the flow tended to break down.
Three or four times as much suction was required at
all incidences to make the separated flow readhere.
With blowing, still larger quantities were necessary,
but the spanwise distribution of the flow from the slot
was unsatisfactory. Two different slot shapes were
tested on the model, one with a sharp beak to the
front lip, the other with a rounded entry. Intermit-
tent separation of the flow occurred in each case.
The phenomena may be of a fundamental character
and associated with the profile shape rather than with
the shape of the slot entry.



N-16947C*

Aeronautical Research Council (Gt. Brit.)
CONTRIBUTIONS TO UNSTEADY AEROFOIL
THEORY. IX. A NUMERICAL COMPARISON OF
VARIOUS APPROXIMATE THEORIES FOR THE
OSCILLATING AEROFOIL OF LARGE SPAN.
Katterbach. August 24, 1948. 27p. diagrs., 6 tabs.
(ARC 11,736; ARC 0. 755)

The lift coefficients of an oscillating elliptical wing
for the reduced frequencies Q = ivb/v = 0, 0. 5,
1.0, ... 6. 5, are calculated according to four well-
known approximate theories which deal with the un-
steady motion of an airfoil; the results, expressed
in complex form, are compared with each other and
with strip theory. The theories are described brief-
ly, and the equations required for the calculation are
derived.


N-16967*

Aeronautical Inspection Directorate (Gt. Brit.)
MECHANICAL THUMB HARD DRYING TIME TEST-
ER FOR PAINTS. A. T. Josling. May 1952. 9p.
photos., diagrs. (AID Rept. APP. 2)

An original mechanical thumb drying time tester,
capable of reproducing results with varying types of
paints, mention being made of protection coating and
other metal finishes.


7


N-16968*

Aeronautical Inspection Directorate (Gt. Brit.)
APPARATUS FOR OBTAINING A WIDE RANGE OF
HUMIDITY AND TEMPERATURE CONDITIONS IN A
STREAM OF AIR. A. T. Josling. May 1952. 4p.
diagrs., photos. (AID Rept. APP. 3)

A new type of apparatus for obtaining any desired
humidity in a stream of air is described. It is based
on the principle of saturating air at a suitable tem-
perature and then heating it to the working tempera-
ture. Saturation is obtained by means of a special
splashing device without atomization. Since the sat-
uration temperature to give the desired humidity at
the working temperature, which may be i '.%,1 1,.
within its wide range, is known from ti',,,. metricc
tables, an actual determination of the humidity
achieved is rendered unnecessary. The apparatus is
suitable for blushing tests on paints and dopes, cor-
rosion tests on platings, protective coatings, oils
and greases.


N-169694

Aeronautical Inspection Directorate (Gt. Brit.)
DEVICE FOR COLD BEND TESTS ON PAINT TEST
PANELS. A. T. Josling. May 1952. 3p. diagrs.
(AID Rept. APP. 4)

A device enabling bend tests to be performed at cor-
rect temperatures. A hinged test panel holder is
also shown.


N-16970*

Aeronautical Inspection Directorate (Gt. Brit.)
THE MEASUREMENT OF THE WALL THICKNESS
OF A LONG TUBE. H. R. Davenport and E. W.
Carter. December 5, 1951. 7p. diagrs., photos.
(AID Rept. Metro 4)

The description of a method of measuring the wall
thickness of a long tube by means of an alinement
telescope and auxiliary equipment.



N-17019*

Forest Products Research Lab. (Gt. Brit.)
TRIALS OF TIMBER FOR PLYWOOD MANUFAC-
TURE. PROGRESS REPORT SIXTEEN. June 1952.
9p. 5 tabs. (Forest Products Research Lab.)

Tests were made on three British Guiana species of
wood. They were crabwood (carapa guianensis),
kurokai (protium decandrum), and wamara (swarzia
leiocalycina). It was found that the crabwood would
be unsuitable for British mills but it may be of some
use for conversion in the country of origin if freshly
felled logs are less prone to splitting and shaking.
Kurokai would also be unsuited to British mills but
would be of use in the country or origin. Wamara is
considered unsuitable because of its weight and pro-
cessing difficulties.




Uilivcmot I Vr Jr Li-.if\Lht^f


NACA 3 1262 09079 7472
8 RESEARCH ABSTRACTS NO.


MISCELLANEOUS


NACA Rept. 1037

Errata No. 1 on "GENERAL METHOD AND THER-
MODYNAMIC TABLES FOR COMPUTATION OF
EQUILIBRIUM COMPOSITION AND TEMPERATURE
OF CHEMICAL REACTIONS". Vearl N. Huff,
Sanford Gordon and Virginia E. Morrell. 1951.


NACA TN 2277

Errata No. 1 on "EFFECTS OF COMPRESSIBILITY
ON THE PERFORMANCE OF TWO FULL-SCALE
HELICOPTER ROTORS'. Paul J. Carpenter.
January 1951.


NACA TN 2353

Errata No. 1 on "CHARTS AND TABLES FOR USE
IN CALCULATIONS OF DOWNWASH OF WINGS OF
ARBITRARY PLAN FORM'. Franklin W. Diederich.
May 1951.



UNPUBLISHED PAPERS


N-5292*

FORMS OF WING FLUTTER IN TWO-DIMENSIONAL
COMPRESSIBLE POTENTIAL FLOW. (Flugelsch-
wingungsformen in ebener kompressibler Potential-
strcmung). Nikolaus Rott. July 1952. 44p. diagrs.
(Trans. from Zeitschrift fur angewandte Mathematik
und Physik, v. 1, 1950, p. 380-410)

On the basis of energy considerations, a survey is
given of the possible forms of oscillations for flutter
with two degrees of freedom in a plane compressible
potential flow. It is found that the possible forms do
not much depend on Mach number in the whole range
from 0 to a, with some exceptions in the neighbor-
hood of M = 1, where forms with only one degree of
freedom (torsion) may occur. Conclusions are
drawn for methods of preventing flutter. Limits of
the reduced frequency are given for the possibility of
flutter in two and one degree of freedom. Special
care was given for the case of M = 1.


NACA-Langley 9-16-52 -4000




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