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
Publication Date:
Frequency:
irregular
completely irregular

Subjects

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:
AA00009235:00032

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Full Text

N(


CURRENT NACA REPORTS
NACA Rept. 1135

EQUATIONrS, TABLES, AND CHARTS FOR COM
PRESSIBLE FLOW. Ames Research Staff. 1953.
iii, 69p. diagrs., 25 charts, 2 tabs. (NACA
Rept. 1135. Formerly TN 1428)
This report, which is a revision and extension of TN
1428, presents a compilation of equations, tables, and
charts useful in the analysis of high-speed flow of a
compressible fluid. The equations provide relations
for continuous one-dimensional flow, normal and ob-
lique shock waves, and Prandtl-Meyer expansions for
both perfect and imperfect gases. The tables present
useful dimensionless ratios for continuous one dimen-



shock waves and for cones in a supersonic airstream.,
A second series shows the effects of caloric imper-
fections on continuous one-dimensional flow and on
the flowtr through normal and oblique shock waves.


N~ACA Rept. 1153

ON THE APPLICATION OF TRANSONIC SIMILARITY
RULES TO WINGS OF FINITE SPAN. John R.
Spreiter. 1953. ii, 21p. diagrs. (NACA Rept. 1153.
Formerly 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 TM 1374

KINETIC TREATMENT OF THE NUCLEATION IN
SUPERSATURATED VAPORS. (Kinetische
Behandlung der Keimbildung in ubersattigten
Dampenn. R. Becker and W. During. September
1954. 43p. diagrs. (NACA TM 1374. Trans. from
Annalen der :Physik, Ser. 5, v. 24, 1935, p. 719-752).


* AVAILABLE ON LOAN ONLY
ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW.,
THE REPORT TITLE AND AUTHOR.


The equations o fte individual processes of self
nucleation are utilized through an electrical analogy
to obtain the nucleation frequency. This process is
shown to be shorter and less subject to error than
that of previous investigators since the appearance
of indeterminant integration constants is completely
avoided. With the nucleation frequencies of crystals
and spheres the Ostwald law of stages is reviewed
and modified. In the final section the general
resistance Image Is dlsscussed and mntionln Is made
of the relation of the electrical network and Volmers
formula.


NACA TN 3199



McComb, Jr. September 1954. S4p. diagrs. (NACA
TN 3199)-

Equations are derived for the stress distributions
caused by three types of loading on infinitely long,
circular, semtimonocoque cylinders with flexible
rings. The results are given as formulas for the
stringer loads and shear flows in the shell due to
each type of loading. For each loading case these
formulas can be used to construct tables of influence
coefficients giving stringer loads and shear flows in
the neighborhood of the load due to a unit magnitude
of that: load.


NACA TN 3231
National Advisory Committee for Aeronautics.
BENDING; TESTS ON BOX BEAMS HAVING SOLID-
AND OPEN-CONSTRUCTION WEBS. Aldie E.
Johnson, Jr. August 1954. 25p. diagrs., photos., 2
tabs. (NACA TN 3231)

The results of an exploratory exp~erimnental investiga-
tion of the effects of replacing alternate webs in a
multiweb beam by open, post-stringer construction
are reported. Post-stringer (either upright or in-
clined posts) construction is shown to perform the
function of comparable-weight, solid, fabricated webs
in the stabilization of the compression cover of a
beam in bending both before and after buckling.


NACA TN 3235
National Advisory Committee for Aeronautics.
LOW-SPEED YAWED-ROLLING AND SOME OTHER
ELASTIC CHARACTERISTICS OF TWO 56-INCH-
DIAMETER, 24-PLY-RATING AIRCRAFT TIRES.
Walter B. Horne, Bertrand H. Stephenson and Robert
F. Smiley. August 1954. 108p. diagrs., photos., 6
tabs. (NACA TN 3235)


WASHINGTON 25, D. C, CITING CODE NUMBER ABOVE EACH TITLE;


National Advisory Commit P"r~ eronauties



Research AP acts
O.70 S ER 21, 1954







NACA
RESEARCH ABSTRACTS NO.70


An examination is made of the mechanism of icing
phenomena. Artificial rime deposits of various types
were produced and measurements made to establish
the quantitative dependence of the deposit on temper-
ature, H20-content, wind speed, and material of the
body to be iced. The growth of the rime on metallic
surfaces was observed under the polarizing micro-
scope. An arrangement is described for studying
rime in the open air, the purpose of which was to
discover whether natural icing is a result of tem-
perature effects. In addition, the influence of the
electric field on the type of deposit was examined
and measured.


DECLASSIFIED NACA REPORTS



THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL,
8/23/54

RM L9K23
RM L50B24a
RM L50C21a
RM L50DO4
RM LSOF28


THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM CONFIDENTIAL,
8/31/54

RM A51B14
RM L50J05a
RM L51A08


NACA RM A6G24

THE SUBSONIC AERODYNAMIC CHARACTERISTICS
OF TWO DOUBLE-WEDGE AIRFOIL SECTIONS
SUITABLE FOR SUPERSONIC FLIGHT. Joseph
Solomon and Floyd W. Henney. May 12, 1947. 33p.
diagrs., photos, (NACA RM A6G24) (Declassified
from Confidential, 8/18/54)

High-speed wind tunnel tests have been made to in-
vestigate the aerodynamic characteristics at sub-
sonic speeds of two symmetrical double-wedge air-
foil sections of 4- and 6-percent thickness. Section
coefficients of lift, drag, and quarter-chord pitching
moment are presented for a moderate range of
angles of attack at Mach numbers up to approximate-
ly 0.93. Comparisons are made between the signifi-
cant characteristics of the double-wedge airfoils and
those of the NACA 65-206 airfoils as an index of the
merit of the former at subsonic speeds.


NACA RM A7G15

AN EXPERIMENTAL INVESTIAION AT SUPER-
SONIC SPEEDS OF ANNULAR DUCT INLETS
SITUATED IN A REGION OF APPRECIABLE
BOUNDARY LAYER. Wallace F. Davis, George B.
Brajnikoff, David L. Goldstein and Joseph M.
Spiegel. September 24, 1947. 41p. diagrs., photos.
(NACA RM A7G15) (Declassified from Confidential,
8/18/54)


The low-speed (up to 4 miles per hour) cornering
characteristics of two 56 x 16, type VII, extra-high-
pressure, 24t-ply-rating tires were determined for a
range of vertical loadings, yaw angles, and tire infla-
tion pressures. Locked-wheel drag tests were also
made for one vertical load condition. The quantities
measured included cornering force, drag force, self-
alining torque, pneumatic caster, vertical tire deflec-
tion, rolling radius, and relaxation length. Some
supplementary tests were made which included meas-
urements of tire footprint area, vertical-load-
deflection characteristics, and the variation of tire
radius and width with inflation pressure.



NACA TN 3303

TURBULENT-HEAT-TRANSFER MEASUREMENTS
AT A MACH NUMBER OF 3.03. Maurice J.
Brevoort and Bernard Rashis. September 1954.
21p. diagrs., tab. (NACA TN 3303)

A three-dimensional axially symmetric plug nozzle
was used to obtain flat-plate data on turbulent-heat-
transfer coefficients and recovery factors. The test
results of this paper are for Mach number 3.03 and
for a Reynolds number range of 5.6 x 106 to 6.5 x 107.
The heat-transfer-coefficient results are in good
agreement with theoretical analyses and the
recovery-factor results are in good agreement with
extrapolations of lower Reynolds number data.



MISCELLANEOUS



N-22929

Advisory Group for Aeronautical Research and
Development. DIFFUSION FLAMES IN THE LABO-
RATORY. John Barr. March 1954. 10p. diagrs.
(Advisory Group for Aeronautical Research and
Development. AG11/MI)

A survey of five recent reviews of diffusion flames
is given. All the flames discussed are related to
each other or can be obtained in a similar burner by
varying the fuel type, the fuel and air flows or the
static pressure. This survey tended to emphasize
our lack of knowledge and the need for further re-
search concerning the factors governing the strue-
ture, length and occurrence of diffusion flames'




UNPUBLISHED PAPERS



N-13524*

EXPERIMENTAL INVESTIGATION OF ICING
PHENOMENA. (Experimentelle Untersuchung von
Vereisungserscheinungen) Domenic Melcher.
April 1954. 36ip. diagrs., photos. (Trans. from
Zeitschrift fur angewandte Mathematik und Physik
v. 2, no. 6, 1951, p. 421-443) '







NACA
RESEARCH ABSTRACTS NO.70

The recovery of total pressure after diffusion to a
low subsonic velocity was found to be approximately
two-thirds of that through a normal shock wave oc-
curring at the same free-stream Mach number. The
cause of this low-pressure recovery is the inter-
action between the boundary layer and the back pres-
sure in the diffuser. Compression at a local Mach
number comparable to that of the supersonic stream
will result in large losses in total pressure if the
compression occurs in the presence of an appreci-
able boundary layer.



NACA RM A8A13

PRESSURE RECOVERY AT SUPERSONIC SPEEDS
THROUGH ANNULAR DUCT INLETS SITUATED IN
A REGION OF APPRECIABLE BOUNDARY LAYER.
I ADDITION OF ENERGY TO THE BOUNDARY
LAYER. Wallace F. Davis and George B. Brajaikoff.
April, 1948. 22p. diagrs., photos. (NACA
RM A8A13) (Declassified from Confidential, 8/18/54)

This report contains the results of an experimental
investigation made at Mach numbers between 1. 36
and 2. 01 of the total-pressure recovery attainable
with a model having a nozzle upstream of an annular
duct inlet for the purpose of ejecting high-velocity
air into the boundary layer of the flow along the fore-
body. A comparison of the total-pressure recovery
effective in producing thrust is made between a
hypothetical engine that recirculates air to the intake
and one that does not. It was found that for the
assumed conditions recirculation would increase the
effective recovery about 8 percent.


NACA RM A8B16

AN EXIPERIMENTAL INVESTIGATOR OF NACA
SUBMERGED INLETS AT HIGH SUBSONIC SPEEDS.
I INLETS FORWARD OF THE WING LEADING
EDGE. Charles F. Hall and F. Dorn Barclay.
June 9, 1948. 6p. diagrs., photos. (NACA
RM A8B16) (Declassified from Confidential,
8/24/54)

Results are given of an experimental investigation of
NACA submerged inlets with deflectors on a fighter
airplane model at Mach numbers from 0.30 to 0.875.
Data are presented showing the characteristics of
the model without infets and with inlets 16.7l-percent
wing-root chord forward of the wing leading edge.
Th results indicate that the ram recovery at the
entrance was affected greatly by variations in mass-
flow coefficient and only slightly by Mach number
and angle-of-attack variations.




NACA RM A8C22

INVESTIGATION AT SUPERSONIC SPEED (M = 1. 53)
OF THE PRESSURE DISTRIBUTION OVER A 63o
SWEPT AIRFOIL OF BICONVEX SECTION AT
ZERO LIFT. Charles Wr. Frick and John W. Boyd.
June 10, 1948. 33p. diagrs., photos. (NACA
RM A8C22) (Declassified from Confidential,
8/18/54)


The results of an investigation at supersonic speed
(M = 1. 53) of the distribution of pressure at zero
lift over the surface of a 630 swept airfoil of
biconvex section (7 percent thick) are presented.
The measured pressures ar~e compared with theo-
retical values calculated from thin airfoil theory.



NACA RM A8FO8

PRESSURE RECOVERY AT SUPERSONIC SPEEDS
THROUGH ANNULAR DUCT INLETS SITUATED IN A
REGION OF APPRECIABLE BOUNDARY LAYER.
II EFFECT OF AN OBLIQUE SHOCK WAVE
IMMEDIATELY AHEAD OF THE INLET. George B.
Brajnikf. August 9,1948. 15p. diagrs., photos.
(NACA RM A8F08) (Declassified from Confidential,
8/18/54)

This report contains the results of an experimental
investigation made at Mach numbers between 1.36and
2.01 of the total-pressure recovery attainable with a
model having a ramp ahead of an annular inlet for the
purpose of reducing the entrance Mach number by
means of an oblique shock wave. It was found that
the recovery improved considerably with increasing
ramp angle and that a model with a 150 ramp pro-
duced approximately four-fifths of the recovery
through a normal shock wave occurring at the free-
stream Mach number. Use of ramp angles greater
than 150 resulted in unsteady flow through the induc-
tion system and a reduction in recovery.




NACA RM A8F21

AN EXPERIMENTAL INVSTIGATION AT LAGE
SCALE OF SEVERAL CONFIGURATINS OF AN
NACA SUBMERGED AIR INTAKE. Norman J.
Martin and Curt A. Holzhauser. October 19, 1948.
68p. diagrs., photos., 7 tabs. (NACA RMn A8F21)
(Declassified from Confidential, 8/18/54)
Results of an experimental investigation in the Ames
40- by 80-foot wind tunnel of NACA submerged inlets
installed on a full-scale! model of a fighter airplane
are presented. Tests at various inlet-velocity ratios
and angles of attack indicated the same favorable
characteristics that had been noted at small scale.
A small difference in the magnitude of pressure
recovery measured at full scale and small scale was
mainly accounted for by differences of boundary-
layer thickness. The effect on pressure recovery of
systematic changes of ramp divergence is also
presented.




NACA RM A8F22

INVESIATION AT SUPERSONIC SPEED (MI = 1.53)
OF THE PRESSURE DISTRIBUTOR OVER A 63o
SWEPT AIRFOIL OF BICONVEK SECTION AT SEV-
ERALL ANGLES OF ATTACK. John W. Boyd, Elliott
D. Katzen and Charles W. Frick. September 24,
1948. 41p. diagrs., photos., tab. (NACA RMd A8F22)
(Declassified from Confidential, 8/18/54)







NAGA
RESEARCH ABSTRACTS NO. 70

NACA RM A8L16

INVESTIGATION OF DOWNWASH AND WAKE CHAR-
ACTERISTICS AT A MACH NUMBER OF 1.53. I -
RECTANGULAR WING. Edward W. Perkins and
Thomnas N.Canning. March, 1949. 29p. diagrs.
(NACA RM A8L16) (Declassified from Confidential,
8/18/54)

The results of an experimental investigation of the
downwash and wake characteristics behind a rectan-
gular plan-form wing of aspect ratio 3.5 are pre-
sented. The airfoil section was a 5-percent-thick,
symmetrical double wedge. The! tests were made at
a Mach number of 1,53 and a Reynolds nuber of
1.25 million. A comparison between experimental
and theoretical values of the downwash angles is
made.


NACA RM A9D20

INVESTIGATION OF DOWNWASH AND WAKE CHAR-
ACTERHITICS AT A MACEI NUMBER OF 1.53. II -
TRIANGULAR WING, Edward W. Perkins and
Thomas N. Canning. June 6, 1949. 31p. diagrs.
(NACA RM A9D20) (Declassified from Confidential,
8/18/54)
The results of an experimental investigation of the
downwash and wake characteristics for a triangular
plan-form wing of aspect ratio 2.04 are presented.
The wing had a 5-percent-thick, symmetrical,
double-wedge airfoil section. The-tests were made
at a Mach number of 1.53 and a Reynolds number of
2 million. A comparison between experimental and
theoretical values of (ds/dar)a=0 showed that the
agreement was good within that part of the flow field
where the theory predicts dDWnWRSh for positive
angles of attack of the wing. Within the remainder of
the flow field the agreement was only fair. At finite
angles of attack the measured dDWnWash angles
depart markedly from the values predicted by the
theory.


NACA RM A9F16

A COMPARISON OF TWO SUBMERGED INLETS AT
SUBSONIC AND TRANSONIC SPEEDS. Emmet A.
Mossman. September 15, 1949. 31p. diagrs., photos.
(NACA RM A9F16) (Declassified from Confidential,
8/18/54)

A qualitative excperimental study of the effects of
ramp-wall divergence on ram-pressure recovery
has been made at Mach numbers up to 0.96. It is
shown that the use of ramp-wall divergence consid-
erably extends the Mach number range for satis-
factory submerged-inlet operation. The increase in
Mach number for satisfactory pressure recovery is
attributed to a less severe interaction between the
shock waves and the ramp boundary layer.


NACA RM A9F20

AN EXPERIMENTAL INVESTIGATION AT LARGE
SCALE OF SINGLE AND TWIN NACA SUBMERGED
SIDE INTAKES AT SEVERAL ANGLES OF SIDESLIP
Norman J. Martin and Curt A. Holshauser. August 1,
1949. 31p. diagrs., photo. (NACA RM A9F20)
(Declassified from Confidential, 8/18/54)


The results of an investigation at supersonic speed
(M = 1.53) of the distribution of pressure at angles of
attack over the surface of a 63o swept airfoil of bi-
convex section (7 percent thick) are presented. The
measured pressures are compared with theoretical
values calculated from supersonic lifting-surface
theory.


NACA RM A8116

AERODYNAMIC CHARACTERISTICS AT SUBSONIC
AND SUPERSONIC MACH NUMBERS OF A THIN
TRIANGULAR WING OF ASPECT RATIO 2. I -
MAXIM/UM THICKNESS AT 20 PERCENT OF THE
CHORD. Robert E. Berggren and James L.
Summers. November 19, 1948. 41lp. diagrs., photos.
(NACA RM A8Il6) (Declassified from Confidential,
8/18/54)

This report presents the results of wind-tunnel tests
to determine the variation with Mach nuber, from.
0.50 to 1.49, of the lift, drag, and pitching-moment
characteristics of a 5-percent-thick symmetrical
double-wedge section triangular wing. The results
are compared with results of tests of similar wings
from other sources and with the characteristics cal-
culated by current analytical methods.



NACA RM A8I20

AERODYNAMIC CHARACTERISTICS AT SUBSONIC
AND SUPERSONIC MACH NUMBERS OF A THIN
TRIANGULA WING OF ASPECT RATIO 2. II -
MAXIMUM THICKNESS AT MIDCHORD. Harold J.
Walker and Robert E. Berggren. December 3, 1948.
41p. diagrs., photos. (NACA RM A8I20) (Declassi-
fied from Confidential, 8/18/54)

This report presents the results of wind-tunnel tests
to determine the variation with Mach number between
limits of 0.50 and 1.49 of the lift, drag, and pitching-
moment characteristics of a 5-percent-thick triangu-
lar wing. The results are compared with the char-
acteristics calculated by analytical methods. The
effects upon the characteristics due to change in
location of maximum profile thickness are also
discussed.

NACA RM A8I29

RAM-RECOVERY CHARACTERISTICS OF NACA
SUBMERGED INLETS AT HIGH SUBSONIC SPEEDS*
Charles F. Hall and Joseph L. Frank. November 17>
1948. 44lp. diagrs., photos. (NACA RM A8I29) (De-
classified from Confidential, 8/18/54)

Results are presented of an experimental investiga-
tion of the characteristics of NACA submerged inlets
on a model of a fighter airplane for Mach numbers
from 0.30 to 0.875. The effects on the ram-recovery
ratio at the inlets of Mach number, angle of attack,
boundary-layer thickness on the fuselage, inlet loca-
tion, and boundary-layer deflectors are shown. The
data indicate only a slight decrease in ram-recovery
ratio for the inlets ahead of or just behind the wing
leading edge as Mach number increased, but showed
large decreases at high 1Mach numbers for the inlets
aft of the point of maximum thickness of the wing.








NACA
RESEARCH ABSTRACTS NO. 70
Results of an experimental investigation in the Ames
40- by 80-foot wind tunnel of single and twin NACA
submerged-intake installations on a ful-scale model
of a fighter airplane at several angles of sideslip are
presented. The effect aE sideslip on the pressure-
recovery and flow characteristics is shown. The
effect of sidealip on pressure recovery and duct air-
flow instability was small"

NACA RM A9G15

THE EFFECT OF THE PROPELLER SLIPSTREAM
ON THE CHARACTERISTICS OF SUBMERGED
INLETS. Noel K. Delany. September 9, 1949.
41p. diagrs., photos. (NACA RM A9G15) (Declas-
sified from. Confidential, 8/18/54)

Wind-tunnel tests were made of submerged air inlets
in the fuselage behind the propeller and forward of
the wing of a 1/4l-scale model of a hypothetical
fighter airplane powered by a turbine-propeller unit.
The results are presented for ramps with parallel
and with divergent walls, and show the effect of
propeller operation on the ram-recovery ratio at the
entrance of the duct and at the simulated entrance of
the compressor. The propeller used in these tests
had dual rotation and had eight blades with thin
shanks.


NACA RM A9H11

REDUCTION OF PROFILE DRAG AT SUPERSONIC
VELOCITIES BY THE USE OF AIRFOIL SECTIONS
HAVING A BLUNT TRAILING EDGE. Dean R.
Chapman. November 1, 1949. 31p. diagrs., photo.
(NACA RM A9H11) (Declassified from Confidential,
8/18/54)

A preliminary theoretical and experimental investi-
gation has been made on the aerodynamic character-
istics of blunt-trailing-edge airfoils at supersonic
velocities. The theoretical considerations indicate
that properly designed airfoils with moderately blunt
trailing edges can have less profile drag, greater
lift-curve slope, and a higher maximum. lift-drag
ratio than conventional sections. These predictions
have been substantiated by experimental measure-
ments on airfoils of 10-percent-thickness ratio at
Mach numers of 1. 5 and 2. 0, and at Reynolds
numbers between 0.2 and 1.2 million.


NACA RM AOLO1

WIND-TUNNEL INVESTIGATION AT: MACH NUM-
BERS FROM 0.50 TO 1.20 OF AN ALL-MOVABLE
TRIANGULAR WIING OF ASPECT RATIO 4 ALONE
AND WITH A BODY. Louis S. Stivers, Jr. and
Alexander W. Malick. February 2, 1950. 45p*
diagrs., photos., tab. (NACA RM A9L01) (De-
classified from Confidential, 8/23/54)

Aerodtynamic characteristics from wind-tunnel tests
of semispan models are presented for an all-movable
triangular wing alone and with a body (body attitude
Oo). The wing had an aspect ratio of 4 and had doubly
symmetrical double-wedge sections with maximum
thichness-chord ratios of 0.08 in the streamwise di-
rection. The tests were made at Mach numbers
from 0.50 to about 0.98 and from 1.09 to 1.29. The
corresponding Reynolds nubers varied from about
0.8 x 106 to 1.1 x 106. The experimental results are


compared with calculated characteristics based on
linear theory.


NACA RM A9L29

TESTS OF A SMALL-SCALE NACA SUBMERGED
INLET AT TRANSONIC MACH NUMBERS. L.
Stewart Rolls and George A. Rathert, Jr.
February 23, 1950. 18p. diagrs., photos. (NACA
RM A9L29) (Declassified from Confidential, 8/23/54)

The pressure-recovery characteristics of an NACA
submerged inlet have been measured qualitatively in
the Mach number range 0.60 to 1.08 by the wing-flow
method. High ram-recovery ratios were maintained
up to test-station Mach numbers of 1.03 to 1.08
where, for mass-flow ratios below 0.5, an abrupt
loss in pressure recovery was measured.


NACA RM A50A04

STALL CHARACTERISTICS OBTAINED FROM
FLIGHT 10 OF NORTHIROP X-4 NO. 2 AIRPLANE
(USAF NO. 46-677). Melvin Sadoff and Thomas R.
Sisk. February 27, 1950. 25p. diagrs., photos., tab.
(NACA RM A50AO4) (Declassified from Confidential,
8/23/54)

Some limited results of stall characteristics have
been obtained on a Northrop X-4 airplane. The re-
sults indicated that the motions of the airplane follow-
ing the stall were mild and recovery was effected
rapidly and completely with the use of down-elevon de-
flections. It was shown that the relative mildness of
the stalls was probably due to the flat-top type of lift-
curve characteristic of the 0010-64 airfoil section
used, the plan form, and to the effect of sideslip on
the stalling characteristics. It was further indicated
that the low values of maximum normal-force coeffi-
clents obtained in gradual stall maneuvers could have
been increased if corrective control had been used.


NACA RM A50C13

PRELIMINARY INVESTIGATION OF THE TRAN-
SONIC CHARACTERISTICS OF AN NACA SUB-
MERGED INLET. John A. Axelson and Robert A.
Taylor. June 5, 1950. 44lp. diagrs., photos.
(NACA RM A50C13) (Declassified from Confidential,
8/23/54)

Ram-recovery and pressure-distribution results for
an NACA submerged inlet investigated on a transonic
bump in the Ames 16-foot high-speed wind tunnel are
presented for Mach numbers from 0.70 to 1.15, for
simulated angles of attack of Oo, 4o, and So, and for
mass-flow ratios between 0.25 and 0.68. The results
indicate favorable ram-recovery characteristics for
the inlet over the entire test range of Mach numbers,
with the minimum ram recovery occurring around
1.05 Mach number.



NACA RM A50D27

LONOI[TUDINAL-STABILITY: CHARACTERISTICS OF
THE NORTHROP X-4 AIRPLANE (USAF NO. 46-677).
Melvin Sadoff and Thomas R. Silsk. June 29, 1950.
24p. diagrs., photos., tab. (NACA RM A50D27)
(Declassified from Confidential, 8/23/54)










The results obtained from several recent flights on
the Northrop X-4 No. 2 airplane are presented. In-
formation is included on the longitudinal-stability
characteristics in straight flight over a Mach number
range of 0.38 to about 0.63, the stability characteris-
tics in accelerated flight over a Mach number range
of 0.43 to 0.79, and the short-period longitudinal-
oscillation characteristics at Mach numbers of 0.49
and 0.78. A description and a limited analysis of a
pitch-up which was encountered during the acceler-
ated stability tests at a Mach number of 0.79 and a
normal-force coeffxcient of about 0.45 are also in-
cluded.


NACA RM A50E26

INVESTIGATION AT HIGH SUBSONIC SPEEDS OF
METHODS OF ALLEVIATING THE ADVERSE
INTERFERENCE AT THE ROOT OF A SWEPT-
BACK WING. Lee E. Boddy. August 10, 1950. .31p.
diagrs., photos. (NACA RM A50E26) (Declassified
from Confidential, 8/23/54)

Tests were made of a model having the 50-percent-
chord line of the wing unswept or swept back 350*
The tests with the swept-back wing included modifi-
cations to the body contour and to the wing section in
the region of the wing-body juncture. The divergence
Mach number of the swept-back wing agreed fairly
well with that predicted from tests of the unswept
wing using the simple cosine concepts. Either of the
modifications increased the drag-divergence Mach
number of the swept-back wing about 0.01.


NACA RM A50IO1

SUMMARY REPORT OF RESULTS OBTAINED DUR-
ING DEMONSTRATION TESTS OF THE NORTHROP
X-4 AIRPLANES. Melvin Sadoff and Thomas R.
Sisk. December 13, 1950. 46~p. diagrs., photos.,
tab. (NACA RM A50101) (Declassified from Confi-
dential, 8/23/54)

The results obtained during the demonstration flight
tests of the Northrop X-4 No. 1 and No. 2 airplanes
are presented. Information is included on the static
and the dynamic longitudinal- and lateral-stability
characteristics, the stalling characteristics, and the
buffet boundary'




NACA RM A50114a

INVESTIGATION OF THE DOWNWASH AND WAKE
BEHIND A TRIANGULAR WING OF ASPECT RATIO
4 AT SUBSONIC AND SUPERSONIC MACH NUMI-
BERS. Harold J. Walker and Louis S. Stivers, Jr.
December 12, 1950. 32p. diagrs. (NACA
RM A50Il4a) (Declassified from Confidential,
8/23/54)

Downwash and wake characteristics for a thin, sym-
metrical, triangular wing are presented for angles
of attack up to 100 and for Mach numbers from 0.50
to 1. 29, corresponding to Reynolds numbers between
0.8 x 100 and 1.1 x106. Dowanwash angles are shown
for locations between the 25- and the 758-percent-
semispan stations at a distance of about 0.8 root-
chord length downstream from the trailing edge of
the wing. The location and thickness of the wake are


NACA
RESEARCH ABSTRACTS NO. 70

given for the 50-percent-semispan station at a dis-
tance of 3.44 root-chord lengths downstream from
the trailing edge. A number of comparisons between
the results from experiment and theory are also
included.


NACA RM A50JO9

LIFT AND MOMENT CHARACTERISTICS AT SUB-
SONIC MACH NUMBERS OF FOUR 10-PERCENT-
THICK AIRFOIL SECTIONS OF VARYING TRAILING-
EDGE THICKNESS. James L. Summers and William
A. Page. December 20, 1950. 32p. diagrs., photos.
( ACA RM A50JO9) (Declassified from Confidential,

The results of a wind-tunnel investigation from 0.3 to
0.9 Mach number of the lift and moment character-
istics of four 10-percent-thick circular-are airfoil
sections having trailing-edge thicknesses of 0, 25,
50, and 100 percent of the maximum thickness are
presented. It was observed that increases in the
trailing-edge thickness resulted in increases in
maximum lift coefficient and lift-curve slope at all
Mach numbers and in increases in lift-divergence
Mach number at all lift coefficients. Kirmin vortex
streets were produced by the thick trailing-edge
sections indicating the presence of fluctuating lifts
and high drags. It was further noted that a splitter
plate at the trailing edge and in the chord plane
eliminated the vortex streets.


NACA RM A50JO9b

WIND-TUNNEL INVESTIGATION AT MACH
NUMBERS FROM 0.50 TO 1.29 OF AN UNSWEPT
TAPERED WING OF ASPECT RATIO 2.67 WITH
LEADING- AND TRAILING-EDGE FLAPS -
TRAILING-EDGE FLAPS DEFLECTED. Louis S.
Stivers, Jr. and Alexander W. Mlalick.
December 13, 1950. 45p. diagrs., photo., 5 tabs.
(NACA RM A50JO9b) (Declassified from Confidential,
8/31/54)
Aerodynamic characteristics are presented for an
unswept wing having an aspect ratio of 2.67, a taper
ratio of 0.5, and full-span, 25-percent-chord, plain,
trailing-edge flaps. Sections of the wing were 8-
percent chord thick from the 25- to the 75l-percent-
chord points tapering to sharp leading and trailing
edges. The data were determined from wind-tunnel
tests of the semispan model for a range of angles of
attack from -3o to 120 and for a range of flap de-
flections from -100 to 600 at Mach numbers from
0.50 to about 0.98 and from 1.09 to 1.29. The cor-
responding Reynolds numbers varied from about
0.94 x 106 to 1.27 x 106.



NACA RM A50K10

WIND-TUNNEL INVESTIGATION AT MACHI NUM-
BERS FROM 0.50 TO 1.29 OF AN UNSWEPT,
TAPERED WING OF ASPECT RATIO 2.67 WITH
LEADING- AND TRAILING-EDGE FLAPS -
LEADING-EDGE FLAPS DEFLECTED. Louis S.
Stivers, Jr. and Alexander W. Malick. February 26,
1951. 37p. diagrs., photo., 5 tabs. (NACA
RM A50K10) (Declassified from Confidential,
8/31/54)







NACA
RESEARCH ABSTRACTS NO. 70


Results of a wind-tunnel investigation are presented
for a semispan model of an unswept wing having an
aspect ratio of 2.67, a taper ratio of 0.5, and full-
span, 25-percent-chord, plain, leading-edge flaps.
Sections of the wing in the streamwise direction were
uniform 8-percent chord thick from the 25- to the
75i-percent-chord points tapering to sharp leading
and trailing edges. The data were obtained for a
range of angles of attack from -3o to 12o and for a
range of flap deflections from -200 to 100 at Mach
nubers from about 0.50 to 0.95 and from 1.09 to
1.29. The corresponding Reynolds numbers varied
from about 0.94 x 106 to about 1.27 x 106.






NACA RM A50K15

THE EFFECTS OF CENTRALLY MOUNTED WING-
TIP TANKS ON THE SUBSONIC AERODYNAMIC
CHARACTERISTICS OF A WING OF ASPECT RATIO
10 WITH 350 OF SWEEPBACK. Bruce E. Tinling and
W. Richard Kolk. February 21, 1951. 4l4p. diagrs.,
photos., tabs. (NACA RM A50K15) (Declassified
from Confidential, 8/31/54)

This report presents results of a wind-tunnel investi-
gation of the effects of centrally mounted wing-tip
tanks on the aerodynamic characteristics of a wing
having an aspect ratio of 10 with 350 of sweepback of
the quarter-chord line. Three tip tanks of equal
volume having fineness ratios of 10, 6.67, and 5 were
tested. The effects of a vane near the tank-wing june-
ture were also investigated. Lift, drag, and pitching-
mome~nt data are presented. The Reynolds number of
the tests was varied from 2,000,000 to 10,000,000 at a
M~ach number of 0.25, and the Mach numberwas varied
from 0.25 to 0.90 at a Reynolds number of 2,000,000,






NACA RM A50K<27

THE EFFECTS OF MACH NUMBER AND REYNOLDS
NUMBER ON THE AERODYNAMIC CHARACT'ERIS-
TICS OF SEVERAL 12-PERCENT-THIICK WINGS
HAVING 350 OF SWEEPBACK AND VARIOUS
AMOUNTS OF CAMBER. Bruce E. Tinling and W.
Richard Kolk. February 23, 1951. 68p. diagrs.,
photo., tab. (NACA RM A50K(27) (Declassified
from Confidential, 8/31/54)

Six semispan model wings with 350 of sweepback were
tested: three of aspect ratio 10, and three of aspect
ratio 5. The thickness distribution of the wing sec-
tions was the same from root to tip and there was no
twist. The streamwise sections of the three wings of
each aspect ratio were the NACA 651A012, the NACA
641A312, and the NACA 641A612. Lift, drag, and
pitching-moment data are presented. The Reynolds
number of the tests was varied from 2,000,000 to
10,000,000 at a Mach number of 0.25, and the Mach
nubehr was Oaidfrom 0.25 to 0.02 at a Reynolds


NACA RM A50K27b

WIND-TUNNEL INVESTIGATION AT MACH NUM-
BERS FROM J0.50 TO 1.29 OF AN UNSWEPT,
TAPERED WING OF ASPECT RATIO 2.67 WITH
LEADING- AND TRAILING-EDGE FLAPS FLAPS
DEFLECTED IN COMBINATION. Louis S. Stivers,
Jr. and Alexander W. Malick. February 26, 1951.
41p. diagrs., photo., 7 tabs. (NACA RM A50K27b)
(Declassified from Confidential, 8/31/54)

Aerodynamic characteristics from wind-tunnel tests
of a semispan model are presented for an unswept
wing having an aspect ratio of 2.67, a taper ratio of
0.5, and full-span, 0.25 chord, plain, leading- and
trailing-edge flaps. Sections of the wing were uni-
form 0.08 chord thick from the 0.25 to the 0.75 chord
points and tapered to sharp leading and trailing edges.
The data are presented for a range of angles of attack
from -3o to 12o and for ranges of leading-edge-flap
deflection from -200 to 100 and of trailing-edge-flap
deflection from Oo to 600. The Mach numbers ranged
from about 0.50 to 0.98 and from 1.00 to 1.29 with
corresponding Reynolds numbers varying from about
0.94 x 106 to 1.27 x 106.


NACA RM A50L12

PRELIMINARY INVESTIGATION OF THE DELAY
OF TURBULENT FLOW SEPARATION BY MEANS
OF WEDGE-SHAPED BODIES. George B.
McCullough, Gerald E. NItaberg and John A. Kelly.
March 1, 1951. 28p. diagrs., photos. (NACA
RM A50L12) (Declassified from Confidential,
8/31/54)

A wind-tunnel investigation of pyramidal, wedge-like
bodies as devices for delaying separation of a tur-
bulent boundary layer was conducted. The flow fields
of individual wedges mounted on a flat plate and of
multiple wedges applied to a two-dimensional airfoil
were studied. Substantial increases of maximum
lift and reductions of drag at high lift coefficients
were realized at the expense of doubling the zero-lift
drag. A brief investigation of vane-type vortex
generators gave comparable increases of maxium
lift with about one-half the incremental drag of the
wedges at zero lift.


NACA RM A51A12

THE TRANSONIC CHARACTERISTICS OF 17
RECTANGULAR, SYMMETRICAL WING MODELS
OF VARYING ASPECT RATIO AND THICKNESS.
Warren H. Nelson and John B. McDevitt. May 10,
1951. 91p. diagrs., photos. (NACA RM A51A12)
(Declassified from Confidential, 8/31/54)

An investigation utilizing the transonic-bump tech-
mique was made to determine the aerodynamic char-
acteristics of transonic Mach numbers of 17 rectan-
gular wings having aspect ratios of 6, 4, 2, and 1,
and NACA 63AOXX sections with thickness-to-chord
ratios of 10, 8, 6, 4, and 2 percent. The Mach
nuber range was 0.4 to 1. 1, corresponding under
the test c ndit an to a Reynolds number srang from

without analysis.






NACA
RESEARCH ABSTRACTS NO.70

NACA RM E7K21

COMPARISON OF HOVERING PERFORMANCE OF
HELICOPTERS POWERED BY JET-PROPULSION
AND RECIPROCATING ENGINES. Virginia L.
Brightwell, Max D. Peters and J. C. Sanders.
June 11, 1948. 39p. diagrs., 3 tabs. (NACA
RM E7K21) (Declassified from Confidential, 8/18/54)

An analysis was made to compare the fuel consump-
tions of three jet engines and one reciprocating
engine used to power helicopters in hovering flight at.
sea level. The calculations showed that the conven-
tional reciprocating engine permitted much longer
hovering time than the jet-propulsion engines inves-
tigated, but because the jet-propulsion engines were
lighter than the reciprocating engine, the jet-
propelled helicopters could lift greater disposable
loads.

NACA RM E8C16

STUDY OF STRESS STATES IN GAS-TURBINE DISK
AS DETERMINED FROM MEASURED OPERATING-
TEMPERATURE DISTRIBUTIONS. J. Elmo Farmer,
M. B. Mrillenson and S. S. Manson. July 21, 1948.
41p. diagrs., photos. (NACA RM E8C16) (Declas-
sified from Confidential, 8/18/54)

Results are presented of an experimental investiga-
tion to determine the temperature distribution in an
aircraft-engine gas-turbine disk. High axial- and
radial-temperature gradients were found to exist
both when the engine was accelerated as rapidly as
possible to maximum speed and power output and
when it was gradually brought to these conditions in
a manner typical of normal service operation. Cal-
culated stresses based on the measured temperature
distributions are presented.



NACA RM E8K24

COMPARISON OF PERFORMANCE: OF AN-F-58
AND AN-F-32 FUELS IN J33-A-23 TURBOJET
ENGINE. H. D. Wilsted and J. C. Armstrong.
June 2, 1949. 33p. photos., diagrs., tab. (NACA
RM E8K24) (Declassified from Confidential,
8/18/54)

Performance of fuels corresponding to specifications
AN-F-58 and AN-F-32 was investigated using a
J33-A-23 turbojet engine. Comparatively, AN-F-58
fuel indicated equal performance in terms of engine
thrust and fuel consumption, improved altitude
starting, improved blowr-out limits, higher but non-
detrimental carbon-deposition rate, and iron oxide
contamination. If contaminating iron oxide can be
excluded from the engine, AN-F-58 fuel is con-
sidered satisfactory.



NACA RM E8K26

EXPERIMENTAL PRESSURE DISTRIBUTIONS OVER
WING TIPS AT MACH NUMBER 1.9. I WING TIP
WITH SUBSONIC LEADING EDGE. James M.
Jagger and Harold Mirels. January 27, 1949. 28p.
diagrs., photo. (NACA RM E8K(26) (Declassified
from Confidential, 8/18/54)


NACA RM AS1Bl6

EXPERIMENTAL DOWNWASH AND WAKE CHAR-
ACTERISTICS AT SUBSONIC AND SUPERSONIC
MACH NUMBERS BEHIND AN UNSWEPT, TAPERED
WING OF ASPECT RATIO 2.67 WITH LEADING-
AND TRAILING-EDGE FLAPS. Harold J. Walker,
Louis S. Stivers, Jr. and Luther Beard, Jr*
April 20, 1951. 43p. diagrs. (NACA RM A1B16)
(Declassified from Confidential, 8/31/54)
Results of a wind-tunnel investigation are presented
which show the effects of flap deflection and Mach
number on the downwash and wake characteristics
at distances of 1.80 and 4.74 mean aerodynamic
chord lengths, respectively, downstream from the
midchord line of a semispan model of an unswept'
tapered wing of aspect ratio 2.67 with full-span,
25-percent-chord, plain, leading- and trailing-edge
flaps. Sections of the wing were 8-percent chord
thick from the 25- to the 75-percent-chord points,
and tapered to sharp leading and trailing edges. The
data were obtained for flap deflections ranging from
-200 to 400 at Mach numbers from 0.50 to 1.29. The
corresponding Reynolds numbers varied from
0.94 x 106 to 1.27 x 106.



NACA RM A51C14

A SEMIEMPIRICAL METHOD FOR CALCULATING
THE PITCHING MOMENT OF BODIES OF REVO_
LUTION AT LOW MACH NUMBERS. Edward J.
Hopkins. May 17, 1951. 27p. diagrs., tab. (NACA
RM A51C14) (Declassified from Confidential,
8/31/54)

A semiempirical method, in which potential theory is
arbitrarily combined with an approximate viscous
theory, for calculating the aerodynamic pitching
moments for bodies of revolution is presented. The
method can also be used for calculating the lift and
drag forces. The calculated and experimental
force and moment characteristics of 15 bodies of
revolution are compared.






NACA RM E7Fl3

INVESTIGATION OF SHOCK DIFFUSERS AT MACH
NUMBER 1.85. III MULTIPLE-SHOCK AND
CURVED-CONTOUR PROJECTING CONES. W. E.
Moeckel and J. F. Connors. August 13, 1947. 23p.
diagrs., photos. (NACA RM E7Fl3) (Declassified
from Confidential, 8/18/54)

The configurations tested were; a cone designed to
produce three oblique shocks ahead of the diffuser
inlet in combination with a straight and curved inlet
section, a cone generated by a parabolic are, also in
combination with a curved and a straight inlet see-
tion; a cone-inlet combination designed by the
method of characteristics to produce an isentropic
entrance flow at an angle of attack of Go; and a 300
single-shock cone in combination with a perforated
inlet section. The effect of angle of attack was also
investigated for the isentropic configuration.







NACA
RESEARCH ABSTRACTS NO. 70

An investigation wras conducted at Mach number of
1.01 to determine gpanwise pressure distribution
over a wing tip in a region influenced by a sharp
subsonic leading edge swept back at 700. Except for
pressure distribution on the top surface in the im-
mediate vicinity of the subsonic leading edge, the
maximum difference between linearized theory and
experimental data was 2-1/2 percent (of free-stream
dynamic pressure) for angles of attack up to 40 and
7 percent for angles of attack up to 8o. Pressures
on the top surface nearest the subsonic edge indicat-
ed local expansions beyond values predicted by
linearized theory.

NACA RM E8L24
National Advisory Committee for Aeronautics.
ALTITUDE PERFORMANCE OF AN-F-58 FUELS IN
J33-A-21 SINGLE COMBUSTOR. Ralph T. Dittrich
and Joseph L. Jackson. April 8, 1949. 23p.
diagrs., 2 tabs. (NACA RM E8L24) (Declassified
from Confidential, 8/18/54)

Three fuels conforming to AN-F-58 specification
were investigated in order to determine the influence
of boiling temperatures and aromatic content on alti-
tude performance in single combustor of a 4600-
pound-thrust turbojet engine. Combustion efficien-
cies of the three AN-F-58 fuels were approximately
equal at each simulated engine condition for altitudes
from 5, 000 to 50, 000 feet, 90-percent normal rated
engine speed, and flight Mach numbers of 0.0 and
0.6.


NACA RM E9DOS

CARBON DEPOSITION FROM AN-F-58 FUELS IN A
J33 SINGLE COMBUSTOR. Jerrold D. Wear and
Howard W. Douglass. June 24, 1949. 27p. diagrs.,
poos. i ials(NAC eRM E9DO6) (Declassified

Effects of change in boiling temperature and aromatic
content of AN-F-58 fuels on carbon deposition and
change in simulated engine operating conditions on
carbon deposition were evaluated in a J33 single
combustor. Effect of carbon deposits in the com.
bustor on altitude operational limits was investi-
gated. Carbon deposition increased with increase in
boiling temperature at constant aromatic content and
with increase in aromatic content of the AN-F-58
fuels. Operational limits occurred at slightly lower
altitudes when the combustor assembly contained
carbon deposits than when it did not.

NACA RM E9H15

DOWlNWVASH IN VORTEX REGION BEHIND
TRAPEZOIDAL-WING TIP AT MKACH NUMBER 1.91.
J. L. Cummings, H. Mirels and L. E. Baughman.
November 10, 1949. 39p. diagrs., photos. (NACA
RM E9H15) (Declassified from Confidential, 8/18/54)

Results of an experimental investigation to deter-
mine downwash and wake characteristics in region of
trailing vortex system behind a trapez~oidal-wing tip
at Mach number 1.91 are! presented. The wing was
cut along inner Mach line from the tip and had a 5-
percent-thick symmetrical diamond cross section.
For small angles of attack, the experimental span.
wise variation of -de/da wheree E is the down-
wash angle and ar is the angle of attack) was gen-
erally similar to the variation predicted by linear-
izied theory. At higher angles of attack, differences


between the experimental and theoretical variations
were attributed to a modification of the spanwise
distribution of shed vorticity and a distortion of the
shed vortex sheet.



NACA RMr E9H17

ANALYTICAL DETERMINATION OF EFFECT OF
WATER INJECTION ON POWER OUTPUT OF
TURBINE-PROPELLER ENGINE. Albert O. Ross
and Merle C. Huppert. November 3, 1949. 29p.
diagrs. (NACA RM E9H17) (Declassified from
Confidential, 8/18/54)
An analysis is presented to show the effect of evap-
orative cooling of the charge air during compression
on the performance of a turbine-propeller engine in-
corporating a centrifugal compressor. Calculations
were made `with water as the cooling agent for com-
pressor tip speeds of 1200, 1500, and 1800 feet per
second. Results indicated that a power augmentation
of 200 percent is possible at a compressor tip speed
of 1800 feet per second if sufficient water is evap-
orated during compression to saturate the air at the
compressor outlet.




NACA RMn E9I22a

EXPERIMENTAL PRESSURE: DISTRIBUTIONS OVER
WING TIPS AT MACH NUMER 1.9. II WING TIP
WITH SUBSONIC TRAILING EDGE. Harold Mirels
ad JmepshM Jag~ger.A DeM ember 21(D c941. as
from Confidential, 8/18/54)

An investigation was conducted at Mach number 1.90
to determine experimental pressure distribution over
a wing tip in region influenced by sharp subsonic
trailing edge. Experimental pressure distribution in
region influenced by subsonic trailing edge was gen-
erally in poor agreement with linearized theory.
Difference between theory and experiment was at-
tributed to separation associated with adverse pres-
sure gradient predicted by linearized theory for this
region,


NACA RM E50D05

EXPERIMENTAL INVESTIGATION OF SUPERSONIC
FLOW WITH DETACHED SHOCK WAVES FOR
MACH NUMBERS BETWEEN 1.8 AND 2.9. W. E.
Moeckel. July 5, 1950. 56p. diagrs., photos., 4
tabs. (NACA RM E50D05) (Declassified from
Confidential, 8/23/54)

An experimental investigation of the flow near the
nose of plane and axially symmetric bodies was con-
ducted to check the predictions of theory. The loca-
tion of the detached wave was determined for a
variety of nose forms over a range of Mach numers
from 1.8 to 2.9. At a M~ach number of 1.0, the form
of the detached wave and the pressure distribution
near the nose were also investigated. In addition,
the relation between shock location and flow spillage
was determined for several nose inlets. Most of the
experimental results agreed to good approximation
with the results predicted by the one-dimensional
continuity method.











NACA RM E50D18

EFFECT OF INLET TEMPERATURE AND HUMID-
ITY ON THRUST AUGMENTATION OF TURBOJET
ENGINE BY COMPRESSOR-INLET INJECTION.
Thomas B. Shillito and James L. Harp, Jr. July 3,
1950. 46p. diagrs., photos. (NACA RM E50D19)
(Declassified from Confidential, 8/23/54)

An investigation was conducted to determine the
effect of inlet temperature and humidity on turbojet-
engine performance with water-alcohol injection at
the compressor inlet. At a constant total liquid
flow (fuel flow plus injected flow) the augmented
thrust ratio increased with compressor-inlet temper-
ature. At a given inlet temperature and injection
rate, the thrust, decreased slightly with increasing
humidity*






NACA RM E50EO3a

EXPERIMENTAL INVESTIGATION OF SPREADING
CHARACTERISTICS OF CHOKED JETS EXPANDING
INTO QUIESCENT AIR. Morris D. Rousso and Fred
D. Kochendorfer. August 9,1950. 39p. diagrs.,
photos., tab. (NACA RM E50EO3a. Now RM E51F18)
(Declassified from Confidential, 8/23/54)

The boundaries of single and twin jets discharging in-
to quiescent air are presented for several values of
nozzle-outlet pressure ratio and spacing between no,
sles. The effects of jet temperature, Reynolds num
ber, and humidity on jet spreading are briefly evalu-
ated. The results indicate that for a jet temperature
ratio of 2.6 the pressure boundaries are slightly
smaller than those of corresponding unheated jets
and that the effects of Reynolds number and humidity
are negligible.






NACA RM E50F12

INVESTIGATION OF PERFORATED CONVERGENT-
DIVERGENT DIFFUSERS WITH INITIAL BOUNDARY
LAYER. Maynard I. Weinstein. August 15, 1950.
26p. diagrs., photo. (NACA RM E50Fl2) (Declas-
sified from Confidential, 8/23/54)
An experimental investigation was made at Mach
number 1.90 of the performance of a series of per_
forated convergent-divergent supersonic diffusers
operating with initial boundary layer, which was
induced and controlled by lengths of cylindrical in-
lets affixed to the diffusers. Supercritical mass-
flow and peak total-pressure recoveries were de_
creased slightly by use of the longest inlets (4 inlet
diameters in length). Combinations of cylindrical
inlets, perforated diffusers, and a subsonic diffuser
were evaluated as simulated wind tunnels having
second throats. Comparisons with noncontracted


NACA
RESEARCH ABSTRACTS NO. 70


NACA RM E50H10

DOWNWASH IN VORTEX. REGION BEHIIND RECTAN-
GULAR HALF-WING AT MACH NUMBER 1.01. John
L. Cummings and Rudolph C. Hfaefeli. October 26,
1950. 43p. diagrs., photos., tab. (NACA
RM E50HI10) (Declassified from Confidential,
8/23/54)

Results of an experimental investigation to deter-
mine downwash and wake characteristics in region of
trailing vortex system behind a rectangular half-wing
at Mach number 1.91 are presented. The wing had a
5-percent thick symmetric diamond cross section
D~eveled to a knife edge at the tip. At small angles of
attack, downwash angles were in close agreement
with predictions of linearized theory based on the
assumption of an undistorted vortex sheet. At higher
angles of attack, the flow was greatly influenced by
the rolling up of the vortex sheet.



NACA RM E50H28

COMBUSTION EFFICIENCY AND ALTITUDE OPER-
ATIONAL LIMITS OF THREE LIQUID HYDRO-
CARBON FUEIS HAVING HIGH VOLUMETRIC
ENERGY CONTENT IN A J33 SINGLE COMBUSTOR.
Edward G. Stricker. November 6, 1950. 22p.
diagrs., tab. (NACA RM E50H28) (Declassified
from Confidential, 8/23/54)
Combustion efficiency and altitude operational limits
were determined in a J33 single combustor for
AN-F-58 fuel and three liquid hydrocarbon fuels
having high volumetric energy content (decalin,
tetralin, and monomethylnaphthalene) at simulated
altitude and combustor inlet-air conditions. At the
conditions investigated, the combustion efficiency
for the four fuels generally decreased with an in-
crease in volumetric energy content. The altitude
operational limits for decalin and tetralin fuels were
higher than for AN-F-58 fuel; monomethylnaphtalene
fuel gave the lowest altitude operational limit.



NACA RM E50H29

INVESTIGATION OF POWER REQUIREMENTS FOR
ICE PREVENTION AND CYCLICAL DE-ICING OF
INLET GUIDE VANES WITH INTERNAL ELECTRIC
HEATERS. Uwe von Glahn and Robert E. Blatz.
December, 1950. 49p. diagrs., photos. (NACA
RM E50H29) (Declassifiedfrom Confidential,
8/23/54)
An investigation was conducted to determine the eleo-
tric power requirements for turbojet-engine inlet
guide vanes with continuous heating and with cyclical
de-icing for a range of icing conditions. Minimum
total power requirements for continuous heating and
dyclical de-icing are presented in terms of average
surface datum temperature. An analysis is included
to extend the experimentally obtained continuous heat-
ing data to vane sizes and icing conditions other than
those investigated. Cyclical de-icing provides a total
power saving as high as 79 percent over continuous







NACA
RESEARCH ABSTRACTS NO.70

NACA RM E50108

ICING CHARACTERISTICS AND ANTI-ICING HEAT
REQUIREMENTS FOR HOLLOW AND INTERNALLY
MODIFIED GAS-HEATED INLET GUIDE VANES.
Vernon H. Gray and Dean T. Bowden. December 5,
1950. 49gp. diagrs., photos. (NACA RM50I08)
(D~eclassified from Confidential, 8/23/54)

Gas temperatures and flow rates required for anti-
icing a two-dimensional cascade of turbojet inlet
guide vanes were determined for hollow and internal-
ly modified vanes. The pressure losses caused by
icing on unheated guide vanes were also determined*
Less heat was required for anti-icing internally
modified blades than for the hollow blades. Pres-
sure losses across the cascade were greater at an
inlet temperature of 22o F than at Oo F because of
the characteristic shapes of ice deposits at the two
temperatures.


NACA RM E50K14

SUPERSONIC TUNNEL INVESTIGATION BY MEANS
OF INCLINED-PLATE TECHNIQUE TO DETERMINE
PERFORMANCE OF SEVERAL NOSE INLETS OVER
MACH NUMBER RANGE OF 1.72 TO 2.18. Jeromne
L. Fox. February 14, 1951. 27lp. diagrs., photos*
(NACA RM E50K14) (Declassified from Confidential,
8/31/54)

A suspended flat plate was used to continuously vary
the Mach number in the NACA Lewis 18- by 18-inch
Mach number 1.91 supersonic tunnel. Useful range
of the tunnl was extended over range of Mach num-
bers from 1.72 to 2.18. Maximum variations in Mach
number of the flow produced at vicinity of nose inlets
at zero angle of attack were 1f0.01 and flow angulari-
ties were less than approximately 0.350. The tech-
nique was applied to determination of pressure re-
covery and mass-flow characteristics of four super-
somec nose inlets over Mach number range produced,


NACA RM E50LO8

INVESTIGATION OF THREE TYPES OF SUPERSONIC
DIFFUSER OVER A RANGE OF MACH NUMBERS
FROM 1.75 TO 2.74. L. Eugene Baughman and
L'awrence I.Gould. March 12, 1951. 37p. diagrs.,
photos. (NACA RM E0L8) (Declassified from
Confidential, 8/31/54)

An investigation was conducted to determine the off-
design pressure recovery, mass-flow, and thrust
characteristics of a normal-shock diffuser, two
single-shock spike diffusers, and three convergent-
divergent perforated diffusers operating in the Mach
nuber range between 1.75 and 2.74'



NACA RM L6ll

DRAG MEASUREMENTS OF A 34o SWEPT-
FORWARD AND SWEPT-BACK NACA 65-009 AIR-
FOIL OF ASPECT RATIO 2.7 AS DETERMINED BY
FLIGHT TESTS AT SUPERSONIC SPEEDS. Sidney
R. Alexander. February 20, 1947. 11p. diagrs.,
photos. (NACA RM L6Ill) (Declassified from
Confidential, 8/18/54)


The data were obtained by tracking rocket-propelled
winged bodies moving at supersonic speeds. The
test results show that for the comparable Mach num-
ber range investigated (M = 0.9 1.30) both the 34o
swept-forward and swept-back airfoils produced
lower values of zero-lift drag than the unswept air-
foil. At Mach numbers between 1.0 and 1.3, the drag
of the swept-back wing was about 50 percent and that
of the swept-forward wing about 65 percent of the
unswept wing,

NACA RM L6K21

LONGITUDINAL STABILITY AND CONTROL CHA-
ACTERISTICS OF A SEMISPAN AIRPLANE MODEL
WITH A SWEPT-BACK TAIL FROM TESTS AT:
TRANSONIC SPEEDS BY THE NACA WING-FLOW
METHOD. John A. Zaloveik and Richard H. Sawyer.
March 28, 1947. 30p. diagrs., photos., tab. (NACA
RM L6K21) (Declassified from Confidential, 8/18/54)

The model was mounted in such a way as to permit it
to assume a position of zero pitching moment about
the center of gravity at 27 percent of the mean aero-
dynamic chord. Measurements were made of lift
and angle of attack for trim for several stabilizer
and elevator settings. Because of the chordwise
variation of Mach number in the test region, the
effective Mach number for the wing of the model was
lower than that for the tail of the model.

NACA RM L6L09

OBSERVATIONS ON AN AIL~ERON-FLUTTER IN-
STABILITY ENCOUNTERED ON A 450 SWEPT-
BACK WING IN TRANSONIC AND SUPERSONIC
FLIGHT. Marvin Pitkin, William N. Gardner and
Howard J. Curfman, Jr. April 11, 1947. 23p.
diagrs., photos. (NACA RM L6L09) (Declassified
from Confidential, 8/18/54)

Flight data are presented showing the free-floating
and oscillatory characteristics of the wing and
aileron. It is shown that aileron vibration occurred
at flight velocities corresponding to a Mach number
range from 1.03 to 1.4. The frequencies of the
aileron and wing oscillations were identical at a
given Mach number, were of the order of 100 cycles
per second, and increased with Mach number. The
test data indicate that aileron compressibility flutter
is delayed in appearance but not prevented by swreep-
back.


NACA RM L6L24

DRAG OF A WING-BODY CONFIGURATION CON-
SISTING OF A SWEPT-FORWARD TAPERED WING
MOUNTED ON A BODY OF FINENESS RATIO 12
MEASURED DURING FREE FALL AT TRANSONIC
SPEEDS. Jim Rogers Thompson and Charles W.
Mathews. March 13, 1947. 15ip.diagrs., photos.
(NACA RM L6L24) (Declassified from Confidential,
8/18/54)

The total drag and the drag of the wing were meas-
ured separately. These measurements were made
to determine optimum aerodynamic shapes and con-
figurations for use in the transonic and supersonic
velocity ranges, and show that the drag of the com-
plete configuration rose almost linearly from 0.07 of
atmospheric pressure per unit of frontal area at a
Mach number of 0.90 to 0.30 of atmospheric pressure
at a Mach number of 1. 02.







NACA
RESEARCH ABSTRACTS NO.70

At a constant leading-edge sweep of 450 no orderly
variation of drag coefficient with taper ratio occurs,
the variation being dependent upon the Mach number.
Maximum thickness, leading-edge, and trailing-edge
sweep are all important in determining the drag co-
efficient of a tapered wing. A comparison is made
between the results of theoretical drag calculations
of tapered wings and applicable experimental values
derived herein.



NACA RM L7GO2

A TORSIONAL STIFFNESS CRITERION FOR PRE-
VENTING FLUTTER OF WINGS OF SUPERSONIC
MISSILES. Bernard Budiansky, Joseph N. Kotanchik
and Patrick T. Chiarito. August 28, 1947. 14p.
diagrs., tab. (NACA RM L7GO2) (Declassified from
Confidential, 8/18/54)

Failures probably due to flutter were encountered in
NACA flight tests of several rocket-powered, drag-
research missiles that were intended to attain Mach
numbers of about 1.4. The wing failures of these
missiles led to the development of a simple, semi-
rational torsional stiffness criterion for preventing
flutter of uniform, sweptback or unswept missile
wings that attain supersonic speeds. Results of
missile flights at speeds up to Mach number 1.4
demonstrate the usefulness of the formula.



NACA RM L7K20

WIND-TUNNEL INVESTIGATION OF A WING-
FUSELAGE COMBINATION WITH EXTERNAL
STORES. H. Norman Silvers and Kenneth P.
Spreemann. July 9, 1948. 55p. diagrs. (NACA
RM L7K20) (Declassified from Confidential,
8/18/54)
Contains drag measurements of several external-
store installations specifically designed to delay the
advent of airplane buffet due to stores which has been
recognized in wing-tunnel data as a sharp increase
of the drag coefficient of the external-store installa-
din.Resulits indi ae tat rtihe hrrnume may
be obtained by locating the store maximum thickness
at large chordwise distance from the wing maximum
thickness or by attaching the store flush to the wing
lower surface, thus eliminating the pylon suspension
system typical of current installations.



NACA RM L8A28a

FACTORS AFF CIG 1AnTERA STABILITY AND

Thomas A.Toll. May 13, 1948. 19p. diagrs.
(NACA RM L8A28a) (Declassified from Confidential,
8/18/54)

The effects on dynamic lateral stability and control-
lability of some of the important aerodynamic and
mass characteristics are discussed and methods are
presented for estimating the various stability param-
eters to be used in the calculation of the dynamic
lateral stability of airplanes with swept and low-
aspect-ratio wings.


NACA RM L6L30

DRAG MEASUREMENTS OF A SWEPT-BACK WING
HAVING INVERSE TAPER AS DETERMINED BY
FLIGHT TESTS AT SUPERSONIC SPEEDS. Sidney
R. Alexander. April 8, 1947. 12p. diagrs., photo.
(NACA RM L6L30) (Declassified from Confidential,
8/18/54)

The test results showed that for the comparable
Mach number range investigated (M = 1.0 to 1.275),
the tapered wing produced values of drag coefficient
that averaged about 30 percent lower than those of
the 340 swept-back untapered wing and about 20 per-
cent higher than those of the 450 swept-back untaper-
ed wing. At Mach numbers of 1.1 and 1.2, the
tapered wing revealed drag coefficients of 0.0195 and
0.0225, the latter value being the maximum value
obtained for this arrangement. The data were ob-
tained by radar tracking of the rocket-propelled
winged body moving at supersonic speeds.

NACA RM L7A03

FORCE AND LONGITUDINAL CONTROL CHARAC-
TERISTICS OF A 1/16 SCALE MODEL OF THE
BELL XS-1 TRANSONIC RESEARCH AIRPLANE AT
HIGH MACH NUMBERS. Axel T. Mattson. May 21,
1947. 32p. diagrs., photo., tab. (NACA RM L7AO3)
(Declassified from Confidential, 8/18/54)
The results given do not present completely the force
and longitudinal control characteristics of the model.
General trends are illustrated, however, which can
be qualitatively analyzed for level-flight Mach num-
bers up to 0.93. A large increase in drag coefficient
occurs beyond a Mach number of 0.78. At a Mach
number of approximately 0.825 an initial lift force
break occurs. This force break, up to a Mach num-
ber of approximately 0.875, is not severe, although
elevator-control effectiveness is decreasing. At a
Mach number of 0.9, use of the stabilizer as a trim
control is required.

NACA RM L7E08

MEASUREMENTS OF THE EFFECTS OF THICK-
NESS RATIO AND ASPECT RATIO ON THE DRAG
OF RECTANGULAR-PLAN-FORM AIRFOILS AT
TRANSsONIC SPEEDS. Jm Roger sThom~psodi ads,

photo. (NACA RM L7EO8) (Declassified from
Confidential, 8/18/54)

rf 1hepresen paper rresults are pre unted fo0 to
sections and aspect ratios of 7.6 and 5.1) and are
compared with results for other airfoils of the series
which were previously reported. At sonic and low
supersonic speeds the pressure-drag coefficient was
'""::u les to vaou ie::"Yeation otesur fthe th0 k-
ands 01 bti between values of thickness ratio of 0.06
and 0.09 the exponent was somewhat less than 2.

NACA RM L7E26

FLIGHT TESTS TO DETERMINE THE EFFECT OF
TAPER ON THE ZERO-LIFT DRAG OF WINGS AT
LOW SUPERSONIC SPEEDS. Sidney R. Alexander
and Robert L. Nelson. July 13, 1947. 19p. diagrs*,
photos. (NACA RM L7E26) (Declassified from
Confidential, 8/18/54)








NACA
RESEARCH ABSTRACTS NO.70

NACA RM L8H31

AN INVESTIGATION OF THE LOW-SPEED STATIC
STABILITY CHARACTERISTICS OF COMPLETE
MODELS HAVING SWEPTBACK AND SWEPT FOR-
WARD WINGS. M. Leroy Spearman and Paul
Comisarow. November 19, 1948. 51p. diagrs., tab*
photos. (NACA RM L8H31) (Declassified from
Confidential, 8/18/54)

An investigation has been conducted in the Langley
300 mph 7- by 10-foot tunnel to determine the static
stability characteristics at low speeds of complete
models with various swept wings so that comparisons
might be made with available theoretical and
emphirical methods of predicting the stability charac-
teristics. Longitudinal and lateral stability charac-
teristics, flaps up and down, were obtained for
models having 00, 150, 300, and 450 sweptforward
and sweptback wings*

NACA RM L8J28

SUMMARY OF RESULTS OF TUMBLING INVESTI-
GATIONS MADE IN THE LANGLEY 20-FOOT FREE-
SPINNING TUNNEL, ON 14 DYNAMIC MODELS.
Ralph W. Stone, Jr. and Robert L. Bryant.
December 31, 1948. 91p. diagrs., photos., 23 tabs.
(NACA RM L8J28) (Declassified from Confidential,
8/18/54)

The tumbling characteristics of dynamic models of
14 specific airplane designs tested in the Langley 20-
foot free-spinning tunnel for various loadings and
configurations have been summarized. Both tailless
and conventional airplane configurations were con-
sidered. Sweep angles varied from 150 forward to a
600 delta wing. Aspect ratios varied from 1.27 to
10.6. Consideration has also been given to the prob-
lem of accelerations encountered during a tumble,
pilot escape, and use of wing-tip parachutes for
termination of the tumbling motion.


NACA RM L9A17

PRELIMINARY TANK INVESTIGATION OF THE USE
OF SINGLE MONOPLANE HYDROFOILS FOR HIGH-
SPEED AIRPLANES. Douglas A. King and John A,
Rockett. March 22, 1949. 35p. diagrs., photos., 6
tabs. (NACA RM L9A17) (Declassified from
Confidential, 8/18/54)

Presents results of hydrodynamic take-off and landing
tedsof a odel of aohypo e iah h g-sendnaie ane

khe cene h n vty oInstability oc ured during

ofrgd.The hdodyanani resistance was greater



NACA RM L9A21

AERODYNAMIC CHARACTERISTICS OF A WING
WITH QUARTER-CHORD LINE SWEPT BACK: 45o>
ASPECT RATIO 4, TAPER RATIO 0.6, AND NACA
65A006 AIRFOIL SECTION. TRANSONIC-BUMP
METHOD. Joseph Weil and Kenneth W. Goodson.

8C 4M I 9 19) (Dtcasfe 1 1 rh Cono dential,


This paper presents the results of an investigation
by the transonic-bump method of a wing-fuselage
combination employing a wing with the quarter-chord
line swept back 450, with aspect ratio 4, taper ratio
0.6, and an NACA 65A006 airfoil section. Lift, drag,
pitching moment, and root bending moment were ob-
tained for the isolated wing and wing-body configura-
tions in a Mach number range between 0.6 and 1.18.
Effective downwash angles and dynamic-pressure
characteristics in the region of the tail plane were
also obtained and are presented for a range of tail
heights at one tail length. The effects of two wing-
fence arrangements were also investigated.

NACA RM L9A31

STABILITY AND CONTROL DATA OBTAINED FROM
FIRST FLIGHT OF X-4 AIRPLANE. H~ubert M.
Drake. February 7, 1949. 11p. diagrs., photos.,
tab. (NACA RM L9A31) (Declassified from Con-
fidential, 8 18754)

Results are presented from the first flight of NACA
instrumented Northrop X-4 semitailless research
airplane. Data presented include rudder and eleven
position, and angle of sideslip for various airspeeds
up to 290 miles per hour indicated and 11,000 feet
pressure altitude. The airplane exhibited a slight
degree of static longitudinal instability for certain
conditions.

NACA RM L9BO2

FULL-SCALE INVESTIGATION OF A WING WITH
THE LEADING EDGE SWEPT BACK 47. 50 AND
HAVING CIRCULAR-ARC AND FINITE-TRAILING-
EDGE-THICKNESS AILERONS. Roy H. Lange.
March 11, 1949. 16p. diagrs., photo. (NACA
RM L9BO2) (Declassified from Confidential, 8/18/54)

The results of an investigation in the Langley full-
scale tunnel to determine the aerodynamic character-
istics of a wing with the leading edge swept back
47. 50 and having a 20-percent-chord, 50-percent-
span outboard-flap-type aileron are presented in
this paper. The wing had circular-are airfoil sec-
tions, and the aileron was investigated for the
circular-are contour and for a flat-sided contour
with finite trailing-edge thickness. The aileron ef-
fectiveness was determined and the data are present-
ed for a Reynolds number of 4. 3 x 106 and a Mach
number of about 0. 07.

NACA RM L9BO8

SO E W -TNE EXPETIMENT ON SNL

IN TRE HalGH SUBSONICOSPEE9D R3 aE Sherman

poo., f iads tiNACA8 M )L9B08) (Declassified

Results of tests of three wing models with various
aileron configurations are presented. Density had
little effect on the initial magnitude or initial Mach
number of buzz. Buzz frequency decreased with
decrease of density. Initial buzz Mach number de-
creases with increasing angle of attack. Mass
balancing and changes of spring stiffness of the
ailerons in these tests had little effect on buzz. In-
creasing the aileron mass moment of inertia lowers
dheT b th orqun he ling thecaileron at the wing
test results with two empirical analyses is made.











NACA RM L9B10a

THE EFFECT OF REAR CHINE STRIPS ON THE
TAKE-OFF CHARACTERISTICS OF A HIGH-SPEED
AIRPLANE FITTED WITH NACA HYDRO-SKIS.
John A. Ramsen. March 17, 1949. 7Ip. diagrs.,
photo. (NACA RM L9B10a) (Declassified from
Confidential, 8/18/54)

Results are presented from tank take-off tests of a
dynamic model of a hypothetical high-speed airplane
fitted with NACA hydro-skis and having the trans-
verse curvature of the lower rear portion of the fuse-
lage broken by small longitudinal chine strips. For
the configuration tested, both trim and resistance
were considerably reduced by the addition of the
strips from the speed at which the ski emerged to the
speed at which the rear of the fuselage came clear of
the water. Results indicate that fuselage shape has a
large effect on the take-off characteristics for a
hydro-ski configuration in which the rear of the fuse-
lage acts as a planing surface.




NACA RM L9B23a

MEASUREMENTS OF AERODYNAMIC CHARACTER-
ISTICS OF A 350 SWEPTBACK NACA 65-009 AIR-
FOIL MODEL WITH 1/4-CHORD HORN-BALANCED


Ha r l I. J hsandod or e Br w n A il 18,




wonais 0. 55to 1. 15 Data aet cmared with rmoesuts


obtained from previous tests of an equivalent plain-
flap model. Conclusion is that horn balance appears
satisfactory for M < 0. 95 but will not be useful
between M =0. 95 and M =1. 15.





NACA RM LOB25

AERODYNAMIC CHARACTERISTICS OF A WING
WITH QUARTIER-CHORD LINE SWEPT BACK 350
ASPECT RATIO 4, TAPER RATIO 0. 6, AND NACA
65A006 AIRFOIL SECTION. TRANSONIC-BUMP
METHOD. William C. Sleeman, Jr. and Robert E.
Becht. April 21, 1949. 29p. diagrs., photo., tab.
(NACA RM L9B25) (Declassified from Confidential,
8/18/54)

This paper presents the results of an investigation by
the transonic-bump method of a wing-fuselage com-
bination employing a wing with quarter-chord line
swept back 350, aspect ratio 4, taper ratio 0. 6, and
NACA 65A006 airfoil section. Lift, drag, pitching
moment, and root bending moment were obtained for
the wing-alone and wing-body configurations over a
Mach number range of 0. 6 to 1. 18. Effective down-
wash angles and dynamic-pressure characteristics in
the region of a probable tail location were also
obtained and are presented for a range of tail heights
at one tail length.


NACA
RESEARCH ABSTRACTS NO. 70


NACA RM L9C01

APPARATUS FOR OBTAINING A SUPERSONIC
FLOW OF VERY SHORT DURATION AND SOME
DRAG MEASUREMENTS OBTAINED WITH ITS USE,
John E. Yeates, Jr., F. J. Bailey, Jr. and T. J.
Vogle~wede. July 23, 1951. 23p.diagrs., photos.
(NACA RM LOC01) (Declassified from Confidential,
8/18/54)

Contains description of a vacuum-actuated super-
sonic nozzle and a comparison of drag data obtained
from free-fall tests and models tested using this
equipment. The comparison indicates that the appa-
ratus gives drag values in good agreement with free-
fall, free-flight, and wind-tunnel tests and it should
also give good results for comparative drag studies
for which this equipment was primarily designed,



NACA RM L9C11

STUDY BY NACA WING-FLOW METHOD OF
TRANSONIC DRAG CHARACTERISTICS OF A
BLUNT-NOSE BODY OF REVOLUTION AND COM-
PARISON WITH RESULTS FOR A SHARP-NOSE
BODY. J. Ford Johnston and Mitchell Lopatoff.
April 26, 1949. 26p. diagrs., photos. (NACA
RM L9C11) (Declassified from Confidential,
8/18/54)


Cotissu ieas by NA C wn-l m tod of pres-


vare iatrion in nsonic spesued range. abdy

NACAuio RM t LOC18 -010 oeile tzr


MCH NMBES. ilon D. Haewt rsumpreys.a Mhay 6,e

RMsieato L918 (Dclassifed fromt Conid entaln d

Prssreditributions tests geand sichliren photod-d




raphs1 ofD the01 NAAIRL 6401 and 64A02 iroNis
Mndcated ES thttepimary ffectof inreasin thea
trailng-e dgelangl fed from So o 1oastodecrease



the loading over the rear portion of the airfoil. The
effect was more pronounced with increasing Mach
numbers and lift coefficients.



NACA RM L9C25

ERROR IN AIRSPEED MEASUREMENT DUE TO
STATIC-PRESSURE FIELD AHEAD OF SHARP-
NOSE BODIES OF REVOLUTION AT TRANSONIC
SPEEDS. Edward C. B. Danforth and J. Ford
Johnston. August 19, 1949, 31p. diagrs., photos.
(NACA RM L9C25) (Declassified from Confidential,
8/18/54)








NACA
RESEARCH ABSTRACTS NO.70

Contains measurements of static-pressure error at
several distances ahead of two sharp-nose bodies of
revolution at zero angle of attack in the transonic
speed range by the NACA wing-flow method. By
application of the linearized subsonic theory and the
transonic similarity rule to the data contained here-
in, the position error at any reasonable distance
ahead of any sharp-nose body of revolution with an
approximately parabolic nose profile and fineness
ratio greater than about 4. 5 can be predicted at any
Mach number.


NACA RM. L9D01

PRELIMINARY WIND-TUNNEL INVESTIGATION AT
HIGH-SUBSONIC SPEEDS OF PLANING-TAIL,
BLENDED, AND AIRFOIL-FOREBODY SWEPT
HULLS. John M. Riebe and Richard G. MacLeod.
September 12, 1949. 33p. diagrs., photos., 3 tabs.
(NACA RM L9D01) (Declassified from Confidential,
8/23/54)
A preliminary investigation was made in the Langley
high-speed 7- by 10-foot tunnel to determine the high
subsonic aerodynamic characteristics of three dif-
ferent types of flying-boat hull: namely, a planing-
tail hull, a blended hull, and an airfoil-forebody
swept hull. For comparative purposes a body of
revolution representative of the fuselage of a modern
high-speed airplane was also included. The hull and
fuselage data presented include the forces and mo-
ments of a 51.30 sweptback wing; the models were
tested as reflection-plane half-models on the side
wall of the tunnel. The results include the lift, drag,
and pitching-moment coefficients through angles of
attack from -lo to 4o up to Mach numers to 0.99*


NACA RM: L9DO6

WIND-TUNNEL INVESTIGATION AT HIGH SUB-
SONIC SPEEDS OF THE LATERAL-CONTROL
CHARACTERISTICS OF AN AILERON AND A
STEPPED SPOILER ON A WING WITH LEADING
EDGE SWEPT BACK 51. 30. Leslie E. Schneiter
and John R. Hagerman. June 7, 1949. 34p. diagrs.,
photo. (NACA RM L9D06) (D~eclassified from
Confidential, 8/18/54)

Contains results and discussion of an investigation
to a Mach number of about 0.9 of the lateral-control
characteristics of a 20-percent-chord by 39-percent.
semispan plain sealed aileron and a 60-percent-span
stepped spoiler on a semispan-wing model of aspect
ratio 3.0 and leading edge swept back 51.30. The
general wing aerodynamic characteristics were also
determined in this investigation.


NACA RM LDO6a

MEASUREMENT OF THE DYNAMIC LATERAL
STABILIY OF THE DOUGLAS D-558-1 AIRPLANE
(BUAERO NO. 37071) IN RUDDER KCCIKS AT A MACH
NNMp Rgr072 sHu Art M rake 6Mayle, 1949.
classified from Confidential, 8/18/54)

Contains results of flight measurements of the

dynamic~~~~~~~~~~~ ltr saii o h oga )58 i-


NACA RM L9DO7

INVESTIGATIONS AT SUPERSONIC SPEEDS OF 22
TRIANGULAR WINGS REPRESENTING TWO AIR-
FOIL SECTIONS FOR EACH OF 11 APEX ANGLES.
Eugene S.Love. May 10, 1949. 100p.diagrs.,
photos., 3 tabs. (NACA RM L9D07) (Declassified
from Confidential, 8/18/54)

The results of tests of 22 triangular wings, repre-
senting two leading-edge shapes for each of 11 apex
angles, at Mach numbers of 1.62, 1.92, and 2.40 are
presented and compared with theory. All wings have
a common thickness ratio of 8 percent and a common
location of maximum thickness point of 18 percent.
Lift, drag, and pitching moment are given for all
wings at each Mach number. The relation of transi-
tion in the boundary layer, shocks on the wing sur-
faces, and characteristics of the pressure distribu-
tions is discussed for several wings.


NACA RM L9D11

FLIGHT INVESTIGATION OF THE JETTISONABLE-
NOSE METHOD OF PILOT ESCAPE USING ROCKET-
PROPELLED MODELS. Reginald R. Lundrstrom and
Burke R. O'Kelly. June 2, 1949. 27p. diagrs.,
photos., 2 tabs. (NACA RM L9D11) (Declassified
from Confidential, 8/18/54)

Results of an investigation at a Mach number of 0.87
of the jettisonable-nose method of pilot escape, using
rocket-propelled models, is presented. With a fin-
stablized nose section the accelerations produced are
well within human tolerance. The drag-weight ratios
of nose and rear bodies were such that the decelera-
tion of the nose was less than that of the rear body.
The shielding effect of the nose on the rear body was
appreciable, and forcible separation appears neces-
sary.


NACA RM L9D13

MEASUREMENTS OF AILERON EFFECTIVENESS
OF BELL X-1 AIRPLANE UP TO A MLACH NUMBER
OF 0.82. Hubert M. Drake. June 20, 1949. 7p.
diagrs. (NACA RMV L9D13) (Declassified from
Confidential, 8/18/54)

Contains results of flight measurements of the aller-
on effectiveness of the Bell X-1 airplane up to a
Mach number of 0.82. The data indicate that Mach
number has little effect on the alleron effectiveness
up to a Mach number of 0.82.


NACA RM L9D20

THE EFFECT OF AIR-JET AND STRIP MODIFI-
CATIONS ON THE HYDRODYNAMIC CHARACTER-
ISTICS OF THE STREAMlLINE FUSELAGE OF A
TRANSONIC AIRPLANE. Bernard Weinflash,
K~enneth W. Christopher and Charles L. Shuford, Jr.

eML3,D2 )4D 1 astfe fom Cnfid nial,8/18/54)

Specific free-to-trim tests were made on a 1/12-
size model of a streamline fuselage modified by pat-



ed chine configuration. Tests were also made of a










simulated multiple-step configuration. The effect of
air flow on both the chine and step configurations was
studied. In addition, the effect of substituting nar-
row breaker strips for the rows of jets in the chine
configuration and in three multiple-step configura-
tions was investigated. Data are presented on
resistance, trim, effective hydrodynamic lift, and
spray.



NACA RM L9D20a

MEASURED CHARACTERISTICS OF THIE DOUGLAS
D-558-1 AIRPLANE (BUAERO NO. 37971) IN TWO
LANDINGS. Hubert M. Drake. June 3, 1949. 8p.
diagrs., photos. (NACA RM L9D20a) (Declassified
from Confidential, 8/18/54)

Records were obtained of two landings which indicate
approach and landing speeds of 150 and 115 percent
of the estimated minimum speed. The rates of
descent just prior to the start of the landing flare
were about 1200 to 1800 feet per minute.



NACA RM L9E10

EFFECTS OF MACH NUMBER AND SWEEP ON THE
DAMPING-IN-ROLL CHARACTERISTICS OF WINGS
OF ASPECT RATIO 4. Richard E. Kuhn and Boyd
C. Myers, II. June 27, 1949. 28p. diagrs., photo-
(NACA RM L9E10) (Declassified from Confidential,
8/18/54)
The dampingiin-aon11 charact ristics of hree wings

angles of 3.60, 32.60, and 46.70 at the quarter-chord





tio1 ns alngwif-th ac comarsng with 0 thory.50i





rATTAC FRO M CoT00 AT HIGH SUBONI alr nge 0 +,
SPEDSf. BenadN.Dleyo fetvns and Douglas R. Lord*



NAARM L9E19) Dcasfe rmCnieta,81/4







knLE1) (ess) as locaed trom danfdhe viairiations



momcaentcfi nsa general ipoe hra mri sti fon' the


NACA
RESEARCH ABSTRACTS NO.70

NACA RM L9E25

AERODYNAMIC CHARACTERISTICS OF A WING
WITH QUARTER-CHORD LINE SWEPT BACK 450,
ASPECT RATIO 4, TAPER RATIO 0.3, AND NACA
65A006 AIRFOIL SECTION. TRANSONIC-BUMP
METHOD. Boyd C. Myers, II and Thomas J. King,
Jr. July 20, 1949. 28p. diagrs., photo., tab.
(NACA RM L9E25) (Declassified from Confidential,
8/18/54)

This paper presents the results of an investigation
by the transonic-bump method of a wing-fuselage
combination employing a wing with quarter-chord
line swept back 450 with aspect ratio 4, taper ratio
0.3, and NACA 65A006 airfoil. Lift, drag, pitching
moment, and root bending moment were obtained for
the isolated wing and wing-fuselage configurations
over a Mach number range 0.70 to 1.18. Effective
downwash angles and dynamic-pressure character-
istics in the region of the tail plane were also ob-
taine~d and are presented for a range of tail heights
at one tail length. In order to exrpedite publishing
of these data only a brief analysis has been made.



NACA RM L9F14

EFFECT OF SWEEPBACK ON THE LOW-SPEED
STATIC AND ROLLING STABILITY DERIVATIVES
OF THIN TAPERED WINGS OF ASPECT RATIO 4.
William Letko and Walter D. Wolhart. August 9,
1949. 36p. diagrs., photo. (NACA RM L9F4I)
(Declassified from Confidential, 8/18/54)

Coltin rs1 t of ani stigto aei h
tunnel to determine the effects of sweepback on the







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NACA
RESEARCH ABSTRACTS NO. 70

NACA RM L9F21

RESULTS OBTAINED FROM SECOND FLIGT OF
X-4 AIRPLANE (A.F. NO. 46-676). Walter C.
Williams. July 18, 1949. 13p. diagrs., photos., tab.
(NACA RM. L9F21) (Declassified from Confidential,
8/18/54)

Results are presented from the second flight of the
NACA instrumented Northrop X-4 semitailless re-
search airplane. Data presented include static-
longitudinal-stability measurements and a time
history of a landing approach and landing with the
center of gravity at 19.7 percent mean aerodynamic
chord. The airplane exhibited positive static longi-
tudinal stability and low damping of the lateral
oscillation.


NACA RM L9F28a

AN NACA VANE-TYPE ANGLE-OF-ATI'ACK
INDICATOR FOR USE AT SUBSONIC AND SUPER-
SONIC SPEEDS. Jesse L. Mitchell and Robert F.
Peck. August 16, 1949. 10p. diagrs., photos.
(NACA RM L9F28a) (Declassified from Confidential,
8 18154)

A description is presented of a vane-type angle-of-
attack indicator developed by the NACA for use at
supersonic and subsonic speeds. A brief history of
the development and a wind-tunnel calibration are
given, together with a discussion of the corrections
to be applied to the indicated readings.


NACA RM L9F29

TESTS OF LIFTING SURFACES ON CONICAL AND
CYLINDRICAL PORTIONS OF A BODY AT SUBSONIC
MLACH NUMERS AND AT A MACH NUMBER OF1.2.
Robert 8. Osborne and JTohn B. Wright. September 2,
1949. 22p. diagrs. (NACA RM L9F29) (Declassi-
fied from Confidential, 8/18/54)

Tests have been conducted at a Mach numer of 1.2
and at several subsonic speeds to determine the
possibilities of using a subsonic conical-flow field in
a supersonic free stream to delay the onset of
adverse compressibility effects on lifting surfaces.


NACA RM L9F29a

LATERAL-CONTROL INVESTIGATION OF FLAP-
TYPE CONTROLS ON A WING WITH QUARTER-
CHORD LINE SWEPT BACK 450, ASPECT RATIO 4,
TAPER RATIO 0.6, AND NACA 65A006 AIREDIL
SECTION. TRANSONIC-BUMP METH~OD. Raymond
D. Vogler. August 15, 1949. 22p. diagrs. (NACA
RM L9F29a) (Declassified from Confidential,
8/18/54)

This paper presents the results of a lateral-control
investigation of a 30-percent-chord flap-type control
of various spanwise magnitudes on a wing of 450 of
sweepback of the quarter-chord line, an aspect ratio
of 4, a taper ratio of 0.6, and an NACA 65A006 air-
foil section. Rolling and pitching mromoents and lift
of the semispan wing fuselage were obtained through
a Mach number range of 0.6 to 1.20


NACA RM L9GO6a

INVESTIGATION OF THE NACA 4-(5)(08)-03 TWO-
BLADE PROPELLER AT FORWARD MACH NUM-
BERS TO 0.925. James B. Delano and Melvin M.
Carmel. September 15, 1949. 61p. diagrs., photo.
(NACA RM L9GO6a) (Declassified from Confidential,
8/23/54)

The paper presents propeller characteristics for the
NACA 4-(5)(08)-03 two-blade propeller for forward
Mach numbers up to 0.925 for blade angles from 200
to 7100.


NACA RM L9G20a

RESULTS OBTAINED FROM THIRD FLIGHT OF
NORTHROP X-4 AIRPLANE (A.F. NO. 46-676).
Walter C. Williams. September 9, 1949. 13p.
diagrs., photos., tab. (NACA RM L9G20a) (De-
classified from Confidential, 8/18/54)

Results are presented from the third flight of the
NACA instrumented Northrop X-4 semitailless re-
search airplane. Data presented are steady sidestip
characteristics at 175 miles per hour which show that
directional stability is positive and high and lateral
stability is positive.


NACA RM L9G22

THE STATIC-PRESSUR ERROR OF WING AND
FUSELAGE AIRSPEED INSTALLATIONS OF THE X-1
AIRPLANES IN TRANSONIC FLIGHT. Harold R.
Goodman and Roxanah B. Yancey. July 22, 1949.
20p. diagrs. (NACA RM L9G22) (Declassified
from Confidential, 8/18/54)

Contains measurements of the static-pressure error
of the airspeed installations of the X-1 airplanes ob-
tained from flight tests made in the Mach number
range of 0.8 to 1.32. The static-pressure error was
measured at a point ahead of the fuselage nose and
at a point ahead of the wring tip. Data are presented
showing the variation of the static-pressure error
with Mach number, airplane lift coefficient, and wing
thickness.

NACA RM L9G22a

AERODYNAMrIC CHARACTERISTICS OF A DELTA
WING WITH LEADING EDGE SWEPT BACK 450,
ASPECT RATIO 4, AND) NACA 65A006 AIRFORL
SECTION. TRANSONIC-BUMP METHOD. William
C. Sleeman, Jr. and Robert E. Becht. September 6,
1949. 29p. diagrs., photo., tab. (NACA
RM L9G22a) (Declassified from Confidential,
8/18/54)
This paper presents the results of an investigation
by the transronic-bump method of a wing-fuselage
combination employing a delta wing with leading edge
swept back 450, aspect ratio 4, and an NACA
65A006 airfoil section. Lift, drag, pitching moment,
and root bending moment were obtained for the wing-
alone and wing-fuselage configurations over a Mach
number range of 0.60 to 1.18. Effective downwash
angles and dynamic-pressure characteristics in the
re ion of a probable tail location were also obtained
and are presented for a range of tail heights at one
tail length.









NACA RM L9G25a

STABILITY AND CONTROL DATA OBTAINED FROM
FOURTH AND FIFTH FLIGHTS OF THE NORTHROP
X-4 AIRPLANE (A.F. NO. 46-676). George M.
Valentine. August 4, 1949. 22p. diagrs., photos.,
tab. (NACA RM L9G25a) (Declassified from Con-
fidential, 8/18/54)

Results are presented from the fourth and fifth flights
of the NACA instrumented Northrop X-4 semitailless
research airplane. Data presented from steady level
runs from 140 to 340 miles per hour which showed the
airplane to be longitudinally stable, stick fixed, in the
clean configuration and in the gear-down flaps-
retracted configuration. Data also are presented
from steady sideslip characteristics at 175 to 280
miles per hour which show that the directional sta-
bility is positive and high and lateral stability is
positive.


NACA RM L9G27

AERODYNAMIC CHARACTERISTICS OF A WINGo
WITH QUARTER-CHORD LINE SWEPT BACK 60 *
ASPECT RATIO 4, TAPER RATIO 0.6, AND NACA
65A006 AIRFOIL SECTION. TRANSONIC-BUMP
METHOD. Thomas J. King, Jr. and Boyd C. Myers,
II. September 6, 1949. 32p. diagrs., photos., tab.
(NACA RM L9G27) (Declassified from Confidential,
8/18/54)
The paper presents the results of an investigation by
the transonic-bump method of a wing-fuselage com-
bination employing a wing with quarter-chord line
swept back 60" with aspect ratio 4, taper ratio 0.6,
and 65A006 airfoil. Lift, drag, pitching-moment>
and root-bending moment characteristics were ob-
tained for the isolated-wing and wing-fuselage con-
figurations over a Mach number range from 0.60 to
1.18. In addition, the effect of a fence, mounted at
the mean aerodynamic chord, is also discussed.
Effective downwash angles and dynamic-pressure
characteristics in the region of the tail plane were
also obtained and are presented for a range of tail
heights at one tail length. In order to expedite pub-
liehng of these data only a brief analysis has been


NACA RM L9HO4

INVESTIGATION OF EXTENS[BLE WING-TIP
AILERONS ON AN UNTAPERED SEMISPAN WING
AT Oo AND 450 SWE~EPBACK(. John R. Hagerman
and William M. O'Hare. September 20, 1949. 40p.
di gs;p o d hl, N8A AR LSHO4P) (Declassi-

Contais. rsuls an escu""":: ofe low aped late -

of extensible wing-tip ailerons on each of two un-
tapered wing configurations, one unswept with aspect
ratio of 3.13 and the other swept back 45o with aspect
ratio of 1.59. The results indicate thatr the extensible
ailerons had satisfactory rolling effectiveness at
moderate and high values of lift coefficient but insuf-
ficent rolling effectiveness at low values of lift coef-
ficient for satisfactory application to an airplane.
However, these allerans may be? suitable for use on
some types of missiles. The extensible atlerons
produced yawing moments comparable to those pro-
duced by conventional flap-type ailerons.


NACA
RESEARCH ABSTRACTS NO. 70
NACA RM L9H16

WIND-TUNNEL INVESTIGTION AT LOW TRAN-
SONIC SPEEDS OF THIE EFFECTS OF NUMBER OF
WINGS ON THE LATERAL-CONTROL EFFECTIVE-
NESS OF AN RM-5 TEST VEHICLE. Harold S.
Johnson. November 29, 1949. 15p. diagrs., photo.,
tab. (NACA RM L9H16) (Declassified from
Confidential, 8/18/54)

Contains results of a wind-tunnel investigation con-
ducted to determine the effects of number of wiings
on the rolling characteristics of an RM-5 test
vehicle through a speed range to a Mach number of
0.9.



NACA RM L9H22

AERODYNAMIC CHARACTERISTICS OF A WING
WITH UNSWEPT QUARTER-CHORD LINE, ASPECT
RATIO 4, TAPER RATIO 0.6, AND NACA 65A006
AIRFOIL SECTION. TRANSONIC-BUMP METHOD.
Kenneth W. Goodson and William D. Morrison, Jr.
October 21, 1949. 32p. diagrs., photos., tab.
(NACA RM L9H22) (Declassified from Confidential,
8/18/54)
This paper presents the results of an investigation
by the transonic-bump method of a wing-fuselage
combination employing a wing with unswept quarter-
chord line, aspect ratio 4, taper ratio 0.6, and
NACA 65A006 airfoil section. Lift, drag, pitching
moment, and root bending moment were obtained
for the wing-alone and wing-body configurations
over a Mach number range of 0.6 to 1.18. Effective
downwash angles and dynamic-pressure character-
istics in the region of the tail plane were also ob-
tained and are presented for a range of tail heights
at one tail length.




NACA RM L9H29a

INVESTIGATION OF SOME TURBULENT-
BOUNDARY-LAYER VELOCITY PROFILES AT A
TUNNEL WALL WITHI MACH NUMBERS UP TO 1.2.
Marshall P. T'ulin and Ray Hi. Wright. November 9,
1949. 22p. diagrs. (NACA RM L9H29a) (De-
classified from Confidential, 8/18/54)

Turbulent-boundary-layer profiles at large Reynolds
nubers, of the order of 40,000 based on the momen-
tum thnickness, and teMachh bers up to 1. 2r e

bility on the profiles and on the dislacement and
ryhenun thicknesses are Asc iss The truet




NACA RM L9IO1

A FREE-FLIGHT TECHNIQUE FOR MEASURING
DAMPING IN ROLL BY USE OF ROCKET-POWERIED
MODELS AND SOME INITIAL RESULTS FOR
RECTANGULAR WINGS. James L. Edmondson and
E. Claude Sanders, Jr. Decembelr20, 1949. 25p.
diagrs., photos. (NACA RM L9101) (Declassified
from Confidential, 8/18/54)







NACA
RESEARCH ABSTRACTS NO. 70

A simplified method for obtaining damping in roll
e~xperimnentally at transonic and supersonic speeds
through use of a simple rocket-powered model em-
ploying canted nozzles to apply a torque has been
developed. A description of the! method and results
of the initial flight tests of rectangular wings with
aspect ratio 3.71 are presented.



NACA RM L9106

INVESTIGATION OF THE NACA 4-(3)(08)-03 TWO-
BLADE PROPELLER AT FORWARD MACH NUM-
BERS TO 0.925. James B. Delano and Francis G.
Morgan, Jr. November 2, 1949. 30p. diagrs.,
photos. (NACA RM L9106) (D~eclassified from
Confidential, 8/23/54)

Propeller characteristics are presented for the
NACA 4-(3)(08)-03 blade propeller for forward Mach
numers up to 0.925 for blade angles of 550, 600,
and 650. A comparison of results for this propeller
with those for the NACA 4-(5)(08)-03 propeller is
presented to show the effect of design camber on
propeller performance.



NACA RM L9106a

STATIC STABILIY OF FUSELAGE HAVING A
RELATIVELY FLAT CROSS SECTION. William R'

(NC Rm L90a) ( ecas dfd gro Cn dential,
8/18/54)

Contains results of force tests and flow surveys
made in the Langley free-flight tunnel to determine
the static stability characteristics of several fuse-
lages having a relatively flat cross section and a
high fineness ratio.



NACA RM L9107

INVESTIATION OF THE NACA 4-(4)(06)-04 TWO-
BLADE PROPELLER AT FORWARD MACH. NUM-
BERS TO 0.925. James B. Delano and Daniel E*
Hlarrison. October 28,1949. 39p. diagrs., photos.
(NACA RM L9I07) (Declassified from. Confidential,
8/23/54)

The paper presents propeller characteristics for
the NACA 4-(4)(06)-04 two-blade propeller for for-
ward Machnumbes up to 0.925 for blade angles
from 200 to 700.




NACA RM L9IO8

AERODYNAMIC CHARACTERISTICS OF A WING
WITH QUTARTER-CHORD LINE SWEPT BACK 450,
ASPECT RATIO 6, TAPER RATIO 0.6, AND NACA
65A006 AIRFOIL SECTION. TRANSONIC-BUMP
METHOD. Kenneth W. Goodson and Albert G. Few,
Jr. November 1, 1949. 34p. diagrs., photos., tab.
(NACA RMl L9I08) (Declassified from Confidential,
8/'18/54)


This paper presents the results of an investigation
by the transonic-bump method of a wing-alone and a
wing-fuselage combination employing a wing with
quarter-chord line swept back 450, aspect ratio 6,
taper ratio 0.6, and NACA 65A006 airfoil section.
Lift, drag, pitching moment, and root bending mo-
ment were obtained for the wing-alone and wing-
fuselage configurations over a Mach number range
from 0.60 to 1.18. Effective downwash angles and
dynamic-pressure characteristics in the region of
the tail plane were also obtained and are presented
for a range of tail heights at one tail length.
NACA RM L9128

AN EMPIRICAL CRITERION FOR FIN STABI-
LIZING JETTISONABLE NOSE SECTIONS OF AIR-
PLANES. Stanley H. Scher. December 8,1949.
21p. diagrs., photo., tab. (NACA RM L9128)
(Declassified from Confidential, 8/18/54)

The present paper summarizes the results of inves-
tigations in the Langley 20-foot free-spinning tunnel
which show the inherent instability of models of
jettisonable nose sections. A criterion is presented
from which, for a given center-of-gravity position,
the fin area required for stabilization of airplane
nose sections can be determined.

NACA RM L9JO4

EFFECT OF AIRFOIL SECTION AND TIP TANKS
ON THE AERODYNAMIC CHARACTERISTICS AT
HIGH SUBSONIC SPEEDS OF AN UNSWEPT WING
OF ASPECT RATIO 5. 16 AND TAPER RATIO 0.61.
He NorbmanlSilvr an Kennt P. preh ann b.
(NACA RM L9JO4) (Declassified from Confidential,
8/18/54)
An investigation of a wing of aspect ratio 5.16 and of
taper ratio 0.61 with two airfoil sections and a tip
tank was made in the Langley high-speed 7- by 10-
foot tunnel over a Mach number range that generally
extended from 0.60 to 0.90. In addition to the effect
of airfoil section and tip tank on the lift, drag, and
pitching-moment coefficients of the wing alone and
the wing-tank combination, tests included the effect
of two modifications to the trailing edge of one sec-
tion and horizontal stabilizing fins on the tip tank.
Pitching moment and lift of the tank alone in the
presence of the wing with one airfoil section were
also measured.


NACA RM L9JO5

INVESTIGATION AT MACH NUMBER 1.62 OF THE
PRESSURE DISTRIBUTION OVER A RECTANGULAR
WING WITH SYMMETRICAL CIRCULAR-ARC SEC-
TION AND 30-PERCENT-CHORD TRAILING-EDGE
FLAP. K. R. Czarnecki and James N. Mueller.
January 25, 1950. 81p. diagrs., photos. (NACA
RM L9J05) (Declassified from Confidential,
8/23/54)

Contains the results of an investigation at a Mach
nuber of 1.62 and a Reynolds number range of 0.55
to 1.07 x 106 of the pressure distributions over a
rectangular wing having a 9l-percent-thick circular-
are section and a 30-percent-chord trailing-edge
flap. The characteristics of the air flow at two span-
wise stations, one in the twoe-dimensional flow region
and the other in the region influenced by the airfoil
tip, are discussed.







20


NACA RM L9J13

WIND-TUNNL INVESTIGATION AT LOW SPEED
TO DETERMINE AERODYNAMIC PROPERTIES OF
A JETTISONABLE NOSE SECTION WITH CIRCULAR
CROSS SECTION. Roscoe H. Goodwin. May 19,
1950. 38p. diagrs., photos. (NACA RM L9J13)
(Declassified from Confidential, 8/23/54)

The aerodynamic properties of a model of a jettison-
able nose section with a circular cross section were
determined at low speed from an investigation in the
Langley 20-foot free-spinning tunnel. Force and
moment measurements were made of the nose section
in various positions removed from the fuselage and
in a position simulating its final condition of free fall
(not under the influence of the fuselage). For each
location of the nose, the measurements were obtained
for the nose with and without stabilizing fins attached
for 00 and 50 angle of attack.


NACA RM L9J13a

THE PATH AND MOTION OF SCALE MODELS OF
JETTISONABLE NOSE SECTIONS AT SUPERSONIC
SPEEDS AS DETERMINED FROM AN INVESTIGA-
TION IN THE LANGLEY FREE-FLIGHT APPARA-
TUS. Lawrence J. Gale. May 23, 1950. 35p.
diagrs., photos., 2 tabs. (NACA RM L9J13a) (De-
classified from Confidential, 8/23/54)

An investigation has been conducted on models of two
different designs of jettisonable nose sections where-
in the nose sections have been projected at super-
sonic speeds (Mach number ranged from 1.2 to 1.4)
in the Langley free-flight apparatus. Both nose
designs in the original unstabilized condition turned
away from a nose-first flight attitude and calcula-
tions indicated that a pilot within corresponding full-
scale nose sections would encounter large accelera-
tions (12 negative g for 0.014 sec for one nose
design and 26 negative g for 0.013 sec for the other
design) as a result of this instability. Both nose
designs with fins installed appeared stable, and
calculations indicated that the equivalent motion in a
corresponding full-scale nose section would not
subject the pilot to large accelerations (6 transverse
g).



NACA RM L9719

THE EFFECT OF TIP TANKS ON THE ROLLING
CHARACTERISTICS AT ]HIGH SUBSONIC MACH
NUMBERS OF A WING HAVING AN ASPECT RATIO
OF 3 WITH QUARTER-CHORD LINE SWPT BACK
35o. Richard E. Kuh and Boyd C. Myers, II.
January 17, 1950. 27p. diagrs., photo., 2 tabs.
(NACA RM L9J19) (Declassified from Confidential,
8/23/54)

An investigation of the effect of two wing-tip mounted
tank configurations on the rolling characteristics of
a swept wing through the Mach number range from
0.4 to 0.91 and in the angle-of-attack range from
0.3o to 6.50 was conducted by the free-roll method.
An additional tank configuration was tested through
the Mach number range at a constant angle of attack
of 0.30. Tanks mounted directly on the tip increased
both the aileron effectiveness and the damping in
roll. Pylon-mounted tanks gave a much smaller in-


NACA
RESEARCH ABSTRACTS No. 70

crease in both of these quantities. The variation of
damping in roll with Mach number was similar and
slightly greater than the wing alone for all configura-
tions tested at a constant angle of attack of 0.30.




NACA RM L9J27

MEASUREMENTS OF THE DRAG AND PRESSURE
DISTRIBUTION ON A BODY OF REVOLUTION
THROUGHOUT TRANSITION FROM SUBSONIC TO
SUPERSONIC SPEEDS. Jim Rogers Thompson.
January 16, 1950. 36p. diagrs., photos., 2 tabs.
(NACA RM L9J27) (Declassified from Confidential,
8/23/54)
The drag and pressure distribution on a body of
revolution of fineness ratio 12 were measured by the
free-fall method at large scale under actual flight
conditions throughout a range of Mach number from
0.75 to 1.27. Analysis shows in detail the mecha-
nism of the abrupt drag rise which occurs near the
speed of sound and demonstrates that the theoretical
method of NACA TN 1768 satisfactorily predicts the
shape of the pressure distribution over the test body
at supersonic speeds. Limited information on skin
friction is presented.




NACA RM L9KO2

INVESTIGATION OF EFFECT OF SPAN AND
SPANW1BE LOCATION OF PLAIN AND STEPPED
SPOILER AILERONS ON LATERAL CONTROL
CHARACTERISTICS OF A WING WITH LEADING
EDGE SWEPT BACK 51.30. Jack Fischel and
Alexander D. Hammond. January 18, 1950. 59p.
diagrs., photos. (NACA RM L9KO2) (Declassified
from Confidential, 8/23/54)

Contains results and discussion of a low-speed later-
al control investigation of a 51.30 sweptbackr wiing
equipped with plain and stepped spoiler ailerons hav-
ing various spans and spanwise locations. The
various ailerons were tested on the wing alone and
on the wing with a simulated fuselage, an 0.487-span
drooped nose, an 0.487-span split flap, and with vari-
ous combinations thereof through a large angle-of-
attack range.




NACA RM L9KO3

MAXIMUMI-LIFT INVESTIGATION AT MACH NUM-
BERS FROM 0.05 TO 1.20 OF A WING WIT~H LEAD-
ING EDGE SWEPT BACK 42o. Thomas R. Turner.
February 14, 1950. 21p. diagrs. (NACA
RM L9KO3) (Declassified from Confidential,
8/23/54)

This paper contains the aerodynamic characteristics
in pitch up to maximum lift for a reflection-plane
model having 420 leading-edge sweep, an aspect ratio
of 4, and a taper ratio of 0.625. The Mach number
varied from 0.05 to 1.20 with the Reynolds nuber
varying from 350,000 to 5,000,000. The higher Mach
numbers were obtained by means of the transonic
bump.







NACA
RESEARCH ABSTRAC-TS NO. 70


NACA RM L9KO~4a

A STUDY OF THE DYNAMIC STABILITY OF THE
BELL X-1 RESEARH AIRPLANE. Edward C.
Polhamus. January 10, 1950. 16ip. diagrs., tab.
(NACA RM L9KO4a) (Declassified fromn Confidential,
8 '23/54)

The period and damping of the lateral oscillation of
the Bell X-1 research airplane have been calculated
for a range of Mach numbers and wing loadings for
an altitude of 30,000 feet The effect of changes in
the magnitude of some of the parameters has been
investigated and a comparison with flight results has
been made.




NACA RM L9K11

MEASUREMENTS OF AERODYNAMIIC CHARACTER-
ISTICS OF A 350 SWEPTBACK NACA 65-009 AIR-
FOIL MODEL WITH 1/4-CHORD BEVELLED-
TRAILING-EDGE FLAP AND TRIM TAB BY THE
NACA WING-FLOW METHOD. Harold I. Johnson
and B. Porter Brown. January 6, 1950. 68p'
diagrs., photo. (NACA RM L9K11) (Declassified
from Confidential, 8/23/54)

Contains lift, pitching-moment, and hinge-moment
data from wing-flow tests of a low-aspect-ratio
sweptback airfoil model having a full-span 1/4-chord
bevelled-trailing-edge flap. Mach number range
was 0.55 to 1.15. Conclusion is that bevelled trail-
ing edge is undesirable type of aerodynamic balance
for large speed range because of nonuniform balanc-
ing characteristics at subsonic speeds and loss of
balancing effectiveness at low supersonic speeds.




NACA RM L9K22

RESULTS OBTAINED DURING FLIGHTS 1 TO 6 OF
THE NORTHROP X-4 AIRPLANE (A.F. NO. 46-677).
James T. Matthews, Jr. JTanuary 12, 1950. 19p.
diagrs., photos., tab. (NACA RM L9K22) (Declassi-
fied from Confidential, 8/23/54)

Results are presented for flights 1 to 6 of a NorthroP
X-4 number 2 semitailless research airplane (A.F.
No. 46-677) equipped with NACA instruments. Data
presented include time histories of a complete pull-
up and several short run in level and accelerated
flight, and the effect of dive-brake extension on longi-
tudinal and lateral trim extension.





NACA RM L9C28

AN INVESTIGATION OF THE SPIN, RECOVERY,
AND TUMBLING CHARACTERISTICS OF A 1/20-
SCALE MODEL OF THE NORTHROP X-4 AIRPLANE.
Lawrence J. Gale, Ira P. Jones, Jr. and Jack H.
Wilson. January 4, 1950. 27p. diagrs., photos., 4
tabs. (NACA RM L9K28) (Declassified from
ConfIdential, 8/23/54)


An investigation of the spin, recovery, and tumbling
characteristics of a 1/20-scale model of the
Northrop X-4 airplane has been conducted in the
Langley 20-foot free-spinning tunnel. The effects of
control settings and movements upon the erect and
inverted spin and recovery characteristics of the
model were determined for the model for loading
conditions simulating various degrees of exhaustion
of fuel in the airplane. An investigation was also
made to determine the spin-recovery parachute
requirements.


NACA RM L9LO5

INVESTIGATION OF THE NACA 4I-(4110)61-057-45A
AND NACA 4-(4)(06)-057-45B1 TWO-BLADE SWEPT
PROPELLERS AT FORWARD MACH NUMBERS TO
0.925. James B. Delano and Daniel E. Harrison.
February 6, 1950. 44lp. diagrs., photos. (NACA
RM L9L05) (Declassified from Confidential,
8/23/54)

The paper presents propeller characteristics for the
NACA 4-(4)(06)-057-45A4 and NACA 4-(4)(06)-057-
45B two-blade swept propellers for forward Mach
numbers up to 0.925 for blade angles of 250, 550,
600, 650, and 700. Moderate delays in onset of
adverse compressibility effects were obtained
through use of large amounts of sweep. The meas-
ured delay was approximately 25 percent of the
value predicted by the use of simple infinite-span
sweep theory.



NACA RM L9LO8a

MOTION OF A TRANSONIC AmRPLANE NOSE SEC-
TION WHEN JETTISONED AS DETERMINED FROM
WIND-TUNNEL INVESTIGATIONS ON A 1/25-SCALE
MODEL. Stanley H. Scher and Lawrence J. Gale.
May 26, 1950. 64p. diagrs., photos., tab. (NACA
RMn L9LO8a) (Declassified from Confidential,
8/23/54)

An Investigltion has been conducted writh a 1/25-
scale model in the Langley 300 mph 7- by 10-foot
tunnel, the Langley free-flight tunnel, and the
Langley 20)-foot free-spinning tunnel to determine
the path and motion of a transonic airplane nose sec-
tion when jettisoned. The investigation included
determination of the probable accelerations that
would act on a pilot in the jettisoned nose section.



NACA RM L9L12

PRESSURE DISTRIBUTIONS ON THE BLADE SEC-
TIONS OF THE NACA 10-(3)(066)-033 PROPELLER
UNER OPERATING CONDITIONS. Julian D.
Maynard and Maurice P. Murphy. January 24, 1950.
166p. diagrs., photos., 12 tabs. (NACA RMd L9L12)
(Declassified from Confidential, 8/23/54)

Contains preliminary data obtained by measuring the
pressure distributions on eleven blade sections of
the NACA 10-(3)(066)-033 propeller having section
design lift, coefficients of 0.3 and varying in thick-
ness from about 4 percent at the tip to 16 percent at
the spinner surface. The pressure distributions







22


together with the corresponding section normal-
force coefficients, chordwise-pressure-force coef-
ficients, and pitching-moment coefficients are pre-
sented in the form of tables for a range of angle of
attack from about -2o to 110 and for a Mach number
range from about 0.3 to 1.2.



NACA RM L9L12a

LATERAL-CONTROL INVESTIGATION OF FLAP-
TYPE CONTROLS ON A WING WITH QUARTER-
CHORD LINE SWEPT BACK 350, ASPECT RATIO 4,
TAPER RATIO 0.6, AND NACA 65A006 AIRFOIL
SECTION. TRANSONIC-BUMP METHOD. Robert F.
Thompson. January 25, 1950. 22p. diagrs., tab.
(NACA RM L9L12a) (Declassified from Confidential,
8/23/54)

This paper presents the results of an investigation to
determine the control-effectiveness characteristics
of 30-percent-chord flap-type control surfaces of
various spans on a semispan wing-fuselage model by
the transonic-bump method. The model employed a
wing with the quarter-chord line swept back 350, an
aspect ratio of 4, a taper ratio of 0.6, and an NACA
65A006 airfoil section parallel to the free stream.
Rolling moments, pitching moments, and lift were
obtained through a Mach number range of 0.6 to 1.20.




NACA RM L9L15

THE TIME LAG BETWEEN FLAP DEFLECTION
AND FORCE DEVELOPMENT AT A MACH NUMBER
OF 4. Walter F. Lindsey and Edward F. Ulmann.
February 13, 1950. 11p. diagr., photos. (NACA
RM L9L15) (Declassified from Confidential, 8/23/54)

Presents the results of an investigation at a Mach
number of 4 and a Reynolds number of 5 x 106 of the
time lag between flap deflection and force develop-
ment on a rectangular wing having a 9-percent-thick
symmetrical circular-are profile and a 30-percent-
chord full-span flap. The lag was found to be less
than one-half of a millisecond or less than a 3-1/2
chord movement of the model.




NACA RM L9L16

PRESSURE MEASUREMENTS AT SUPERSONIC
SPEEDS ON A SECTION OF A RECTANGULAR
WING HAVING AN NACA 65-009 PROFILE. Robert
W. Rainey. March 10, 1950. 31p. diagrs., photos.,
tab. (NACA RM L9L16) (Declassified from
Confidential, 8/23/54)

Experimental and predicted pressure distributions
are presented for a section of a rectangular wing
having an NACA 65-009 profile at M = 1.62, 1.93,
and 2.40 at Reynolds numbers of 1.07 x 10 ,
0.97 x 106, and 0.81 x 106, respectively. Integrated
normal-force, pitching-moment, and chordwise co-
efficients, as well as shadowgraphs and a discussion
of the character of flow about the model, are report-
ed. Comparison between the aerodynamic character-
istics of the NACA 65-009 airfoil and a symmetrical
circular-are airfoil of the same thickness ratio is
made.


NACA
RESEARCH ABSTRACTS NO. 70

NACA RM L9L19

FLIGHT INVESTIGATION OF THE EFFECT OF
THICKENING THE AILERON TRAILING EDGE: ON
CONTROL EFFECTIVENESS FOR SWEPTBAICK.
TAPERED WINGS HAVING SHARP- AND ROUND-
NOSE SECTIONS. H. Kurt Strass and Edison M.
Fields. May 2, 1950. 20p. diagrs., photo., tab.
(NACA RM L9L19) (Declassified from Confidential,
8/23/54)

The effect of thickening the trailing edge of an aileron
on control effectiveness and drag at transonic and
supersonic speeds as determined in flight are pre-
sented for a wing having an NACA 0010-64 airfoil
section, a taper ratio of 0.455, an aspect ratio of 3.6,
sweepback of the quarter-chord of 38.10, and a trail-
ing edge angle of 170. The results are compared
with a previous investigation of a similar wing having
a circular-are section. The blunt trailing edge elim-
inated the control reversal obtained with the true-
contour ailerons and greatly improved the control
effectiveness although at some sacrifice in drag.

NACA RM L9L21a

HYDRODYNAMIC FORCE CHARACTERISTICS OF A
STREAMLINE FUSELAGE MODIFIED BY EITHER
BREAKER STRIPS OR ROWS OF AIR JETS SIMU-
LATING CHINES. Bernard Weinflash, Charles L.
Shuford, Jr. and Kenneth W. Christopher.
February 21, 1950. 45~p. diagrs., photos. (NACA
RM L9L21a) (Declassified from Confidential,
8/23/54)

Force tests were made to determine the effect of
trim on the resistance, hydrodynamic lift, and
hydrodynamic moment of a streamline fuselage
modified by either strips or rows of air jets simu-
lating chines. Tests were also made of the model
modified by the strips for three load-on-the-water
conditions and for the model with the longitudinal
curvature of the after half of the fuselage bottom
eliminated.


NACA RM L9L23

PRESSURE DISTRIBUTIONS ON THE BADE SEC-
TIONS OF THE NACA 10-(3)(049)-033 PROPELLER
UNDER OPERATING CONDITIONS. W. H. Gray
and Robert M. Hunt. February 14, 1950. 120p.
diagrs., 11 tabs. (NACA RM L9L23) (Declassified
from Confidential, 8/23/54)

Presents unanalyzed data in tabular form obtained
from pressure-distribution measurements on one of
a family of five related propellers incorporating
16j-series blade sections. Nine radial stations were
investigated with a variation of thickness ratio from
2.6 percent to 8.9 percent and covering a section
Mach number range from 0.375 to 1.197.


NACA RM L9L23a

AN INVESTIGATION OF SEVERAL NACA 1-SERIES
NOSE INLETS WITH AND WITHOUT PROTRUDINGr
CENTRAL BODIES AT HIGH-SUBSONIC MACH
NUMBERS AND AT A MACH NUMBER OF 1.2.
Robert E. Pendley and Harold L. Robinson.
February 21, 1950. 53p. diagrs., photos. (NACA
RM L9L23a) (Declassified from Confidential,
8/23/54)







NACA
RESEARCH ABSTRACTS NO. 70

Measurements of pressure distribution, drag, and
internal-flow pressure loss are presented for three
NACA 1-series note inlets, two of which were fitted
with protruded central bodies. Test M~ach number
and inlet-velocity ratio ranged from 0.4 to 1.2 and
from 0 to 1.34, respectively. Nose-inlet pressure
drag at a Mach number of 1.2, and central-body
effects on subcritical drag, the supercritical drag
rise, and inlet total-pressure loss are discussed.



NACA RM L9L28a

AERODYNAMIC INVESTIGATION AT MACH NUM.
BER 1.92 OF A RECTANGULAR WING AND TAIL
AND BODY CONFIGURATIO AND ITS COMPO-
NENTS. Macon C. Ellis, Jr. and Carl E. Grigsby.
March 1, 1950. 96p. diagrs., photos., 4 tabs.
(NACA RM L9L28a) (Declassified from Confidential,
8/23/54)

An investigation at Mach nuber 1.92 in the L~angley
9-inch supersonic tunnel of a variable body-wing-tail
configuration has been made in order to determine
and to isolate the aerodygnamic effects on each other
of the components of the configuration. The body had
a fineness ratio of 12.5 with a cylindrical midsection
so that the aspect-ratio-4 rectangular wing could be
located at three longitudinal positions along the body.
The variable-incidence-angle rectangular tail was of
the same aspect ratio as the wing, but one-fourth the
wing area, and could be located at three vertical
positions relative to the plane of the wing. The
tests included lift, drag, and pitching-mome~nt meas-
urements of all possible elements and combinations
of this modeL




NACA RM L50A03

LATERAL-CONTROL INVESTIGATION OF FLAP-
TYPE CONTROLS ON A WING WITH UNSWEPT
QUARTER-CHORD LINE, ASPECT RATIO 4,
TAPER RATIO 0.6, AN NACA 65A006 AIRFOR
SECTION. TRANSONIC-BUMP METHOD.
Alexander D. Hammond. March 10, 1950. 20p.
diagrs. (NACA RM L50A03) (D~eclassified from
Confidential, 8/23/54)

This paper presents the results of a lateral-control
investigation of a 30-percent-chord flap-type control
having various spans and spanwise locations on a
wing with unswept quarter-chord line, aspect ratio 4,
taper ratio 0.6, and an NACA 65A006 airfoil section.
Rolling moments, pitching moments, and lift of the
semispan wing-fuselage were obtained through a
Mach number range of 0.7 to 1. 15.




NACA RM L50A03a

AN 8-FQOT AXISYMMETRICAL FIXD NOZZLE
~FOR SUBSONIC MACH NUMBERS UP TO 0.99 AND
FO~R A SUPERSONIC MACH NUMBER OF 1.2.
Virgil S. Ritchie, Ray H. Wright and Marshall P.
Tulin. February 23, 1950. 52p. diagrs., photos.,
2 tabs. (NACA RM L50AWa) (Declassified from
Confidential, 8/23/54)


23


The design and operating characteristics of an 8-
foot-diameter circular nozzle for Mach number 1.2
and for high-subsonic Mach numbers are discussed.
The nozzle was carefully designed and adjusted for
boundary-layer displacement. The flow in the sub-
sonic and supersonic test sections was satisfactorily
uniform for model testing. Consideration was given
to disturbances in the flow, ir regularities at the
wall, humidity effects, and power requirements.



NACA RM L50A12

AERODYNAMIC CHARACTERISTICS OF A WING
WITH QUARTER-CHORD LINE SWEPT BACK 600,
ASPECT RATIO 2, TAPER RATIO 0.6, AND NACA
65A006 AIRFORL SECTION. TRANSONIC-BUMP
METHOD. Boyd C. Myers, 11 and Thomas J. King,
Jr. February 24, 1950. 31p. diagrs., photos., tab.
(NACA RM L50A12) (Declassified from Confidential,
8/23/54)

This paper presents the results of an investigation by
the transonic-bump method of a wing-fuselage combi-
nation employing a wing with quarter-chord line swept
back 600, aspect ratio 2, taper ratio 0.6, and NACA
65A006 airfoil. Lift, drag, pitching-momtent and
root-bending-momnent characteristics were obtained
for the wing-alone and wing-fuselage configurations
over a Mach number range from 0.70 to 1.18. In
addition, effective downwash angles and dynamic-
pressure characteristics in the region of the tail
plane were obtained and are presented for a range of
tail heights at one tail length. In order to expedite
publishing of these data, only a brief analysis has
been made.




NACA RM L50A13

PRELIMINARY INVESTIGATION OF A SUBMERGED
AIR SCOOP UTILIZING BOUNDARY-LAYER
SUCTION TO OBTAIN INCREASED PRESSURE
RECOVERY. Mark R. Nicholas and P. Kenneth
Pierpont. March 17, 1950. 77p. diagrs., photos.,
tab. (NACA RM L50A13) (Declassified from
Confidential, 8/23/54)

Presents results of low-speed tests of a submerged
inlet consisting essentially of a conventional scoop
located in a dimple in the fuselage surface.
Boundary-layer-control systems investigated are
show to provide important increases in perform-
ance. It appears that the flow instability frequently
encountered in the case of twin internally coupled
inlets will be avoided for design high-speed inlet-
velocity ratios as low as 0.5.




NACA RM L50A17

LATERAL-CONTROL INVESTIGATION OF FLAP-
TYPE CONTROLS ON A WING WITH QUARTER-
CHORD LINE SWEPT BACK 600, ASPECT RATIO 4,
TAPER RATIO 0.6, AND NACA 65A006 AIRFORL
SECTION. TRANSONIC-BUMP METHOD. Raymond
D. Vogler. March 2,1950. 21p. diagrs. (NACA
RM L50A17) (Declassified from Confidential,
8/23/54)







24



This paper presents the results of a lateral-control
investigation of a 30-percent-chord flap-type control
of various spanwise magnitudes and locations on a
wing of 600 of sweepback of the quarter-chord line,
an aspect ratio of 4, a taper ratio of 0.6, and an
NACA 65A006 airfoil section. Rolling and pitching
moments and lift of the semispan wing-fuselage were
obtained through a Mach number range of 0.7 to 1.15.



NACA RM L50A18

AERODYNAMIC CHARACTERISTICS OF A WING
WIH UNSWEPT QUARTER-CHORD LINE, ASPECT
RATIO 2, TAPER RATIO 0.78, AND NACA 65A004
AIRFORL SECTION. TRANSONIC-BUMP METHOD.

Marh, 90 e s sia r ~po~t sCtab (ACA
8/23/54)

This paper presents the results of an investigation by
the transonic-bump method of a wing having an un-
swept quarter-chord line, aspect ratio 2, taper ratio
0.78, and NACA 65A004 airfoil. Lift, drag, pitching-
moment, and root-bending-moment characteristics
were obtained over a Mach number range! of from
0.60 to 1.17. Experimental values of lift-curve slope
lateral center of pressure, and drag due to lift are
compared with available theory.




NACA RM L50A23

A DISCUSSION OF THE DESIGN OF HIGHLY SWEPT
PROPELLER BLADES. Richard T. Whitcomb.
May 4, 1950. 31p. diagrs., photos. (NACA
RM L50A23) (Declassified from Confidential,
8/23/54)

A description of the two swept propellers investigat-
ed in the Langley 8-foot high-speed tunnel is pre-
sented, together with the discussions of the numer-
ous assumptions and analyses on which the designs of
these propellers are based.





NACA RM L50A26

PRESSURE DISTRIBUTIONS ON THE BLADE SEC-
TIrONS OF THE NACA 10-(3)(090)-03 PROPELLER
UNDER OPERATING CONDITIONS. Peter J*
Johnson. March 22, 1950. 90p. diagrs., 10tabs.
(NACA RM L50A26) (Declassified from Confidential'
8/23/54)

Presents unanalyzed data in tabular form obtained
from pressure-distribution measurements on one of
a family of five related propellers incorporating
NACA 16-series blade sections. Pressure distribu-
tions were measured at nine radial stations for which
blade-section thickness ratio varied from 0.300 to
0.053 and section Mach number varied from 0.28 to
1.18'


NACA
RESEARCH ABSTRACTS NO.70


NACA RM L50A30

EXPLORATORY INVESTIGrATION OF LEADING-
EDGE CHIORD-EXTENSIONS TO IMPROVE THE
LONGITUDINAL STABILITY CHARACTERISTICS OF
TWO 52o SWEPTBACK; WING;S. G. Che~ster Furlong.
March 10, 1950. 32p. diagrs., photo. (NACA
RM L50A30) (D~eclassified from Confidential,
8/23/54)

Results are presented of exp~loratory tests obtained
with leading-edge wing chord-extensions on two 52o
sweptback wings for the purpose of obtaining satis-
factory longitudinal stability through~ the! lift range.
Both wings exhibit a vortex flow similar to that which
has been observed on triangular wings. One wing of
aspect ratio 2.8 has circular-are airfoil sections,

64-12arfi ec ion T ts wr conduted




NACA RM L50A31

WING-FLOW MEASUREMENTS OF LONGITUDINAL
STABILITY AND CONTROL CHARACTERISTICS OF
A CANARD AIRPLANE CONFIGURATION WITH A
450 SWEPTBACK WING AND A TRIANGULAR ALL-
MOVABLE CONTROL SURFACE. Harold L. Crane
and James J. Adams. August 25, 1950. 53p.
diagrs., photo. (NACA RM L50A31) (Declassified
from Confidential, 8/23/54)

Presented are lift, pitching-moment, rolling-
moment, and drag measurements for a semispan
canard configuration for a Mach number range of
0.55 to 1.14 at Reynolds numbers between 225,000 to
575,000. The configuration consisted of a 450 swept-
back untapered wing and a triangular control surface
mounted on a body of fineness ratio 13.5. The aero-
dynamic center shift with increasing Mach number
was small and the effectiveness of the control sur-
face increased gradually with Mach number with the
result that the longitudinal trim and control char-
acteristics in the transonic speed range at low lift
coefficients were desirable. Also included are
calculated trim curves for various loading conditions
and calculated aeroelastic effects.


NACA RM L50BO2a

INVESTIGATION OF A SIMPLE DEVICE FOR PRE-
VENTING SEPARATION DUE TO SHOCK AND
BOUNDARY-LAYER INTERACTION. Coleman duP.
Donaldson. November 29, 1950. 34p. diagrs.,
photos. (NACA RM L50BO2a) (Declassified from
Confidential, 8/23/54)

Results are presented of a preliminary investigation
of vortex generators introduced into the region of
boundary layer to increase the turbulent mixing, in
an attempt to prevent separation due to boundary-
layer and shock interaction. Prevention of such
separation and some of its adverse effects up to
Mach numbers ahead of the shock of the order of 1.4
appears possible by means of this simple device.
Preliminary flight tests show that this method is ef-
fective in controlling shock-induced separation on
the wing of an airplane at high speed.







NACA
RESEARCH AEISTRACTS NO. 70

NACA RM L50B03a

AERODYNAMIC CHARACTERISTICS OF A WIING
WITH QUA RTER-CHORD LINE SWEPT BACK 450
ASPECT RATIO 6, TAPER RATIO 0.6, AN~D NACA
65A009 AIFO lSE ON. TRANS IIC-BUMP
Morrison, Jr. and Thomas B. Pasteur, Jr. April 6,
1950. 33p. diagrs., photos., tab. (NACA
RM L50BO3a) (Declassified from Confidential,
8/23/54)

This paper presents the results of an investigation
by the transonic-bump method of a wing alone and a
wring-fuselage combination emplo ing a wing writh
quarter-chord line swept back 45 aspect ratio 6,
taper ratio 0.6, and NACA 65A009 airfoil section.
Lift, drag, pitching moment, and root bending
moment were obtained for the wing-alone and wing-
fuselage configurations over a Mach range of 0.60 to
1.18 and through an angle-of -attack range of -4o to
100. Effective downwash angles and dynamic-
pressure characteristics in the region of the tail
plane were also obtained and are presented for a
range of tail heights at one tail length.



NACA RM L50B10

THE DAMIPING IN ROLL OF ROCK(ET-POWERED
TEST VEHICLES HAVING RECTANGULAR WINGS
'W1TH NACA 65-006 AND SYMMETRICAL DOUBLE-
WEDGE AIRFOIL SECTIONS OF ASPECT RATIO 4.5.
Albert E. Dietz and JTames L. Edmnondson. March 29,
1950. 12p. diagrs. (NACA RM L50Bl0) (Declassi-
fled fro Confidential, 8 23 54)

Exp~erimental determination of damping in roH
through a Mach range of 0.85 to 1.45 was con-
ducted for two free-flight models having rectangu-
lar wing plan forms and an aspect ratio of 4.5. One
model had wings with NACA 65-006 airfoil sections
and the other model had wings ~with 6-percent-thick
modified double-wedge airfoil sections.

NACA RM L50B21

PRESSURE DISTRIBUTIONS ON THE BLADE
SECTIONS OF THE NACA 10-(5)(066)-03 PROPEL-
LER UNDER OPERATING CONDITIONS. Albert J.
Evans and Wallace Luchuk. April 18, 1950. 99p.
diagrs., 10 tabs. (NACA RM L50B21) (Declassified
from Confidential. 8723/54)

Presents unanalyzed data in tabular form from pres-
sure distribution measurements on one of a family of
five related propellers incorporating NACA 16-
series sections. Nine radial stations were investi-
gated which had design lift coefficients of 0.50 and
varied in thickness from 16 percent to 4 percent of
the chord. The Mach numer varied from 0.4 to
1.15.


NACA RM L50CO3

PRESSURE DISTRIBUTIONS ON THE BLADE
SECTIONS OF THE NACA 10-(0)(066)-03 PROPEL-
LER UNDER OPERATING CONDUITS. Seymour
Stelnberg and Robert W. Milling. May 18, 1950.
89p. diagrs., 11 tabs. (NACA RM L50COS) (Declas-
sified from Confidential, 8/23/54)


Presents unanalyzed data in tabular form obtained
from pressure-distribution measurements on one of
five related propellers incorporating NACA 16-
series blade sections. Pressure distributions were
mzeasue at ne aii st tion oor0 which e-nde

section Mach number varied fromn 0.35 to 1.14.


NACA RM L50C16

AERODYNAMrIC CHARACTERISTICS OF A WING
WITH UNSWEPT QUARTER-CHORD LINE, ASPECT
RATIO 4, TAPER RATIO 0.6, AND NACA 65A004
AIRFORL SECTION. TRANSONIC-BUMP METHOD.
Boyd C. Myers, II and James W. Wiggins. May 8,
1950. 31p. diagrs., photos., tab. (NACA
RM L50C16) (Declassified from Confidential,
8/23/54)

This paper presents the results of an investigation
by the transonic-bump method of a wing and wing-
fuselage combination employing a wing writh unwept
quarter-chord line, aspect ratio 4, taper ratio 0.6,
and NACA 65A004 airfoil section parallel to free
stream. Lift, drag, pitching-moment, and root-
bending-moment characteristics were obtained for
the two configurations over a Mach numer ranged
from about 0.60 to 1.15. In addition, effective dowEn-
wash angles and dynaic-pressure characteristics
in the region of the tail plane were obtained and pre-
sented for a range of tail heights at one tail length.

NACA RM L50C21

AERODYNAMC CHARACTERISTICS WIrTH FIED3
AND REE TRANSITION OF A MODIFIED DELTA
WIN IN COMBINATION W1T A FUSELAGE AT
HIGH SUBSONI SPEEDS. Edward C. Polhamus and
Thomas J. King, Jr. May 2, 1950. 19p. diagrs.,
photos. (NACA RM L50C21) (Declassfed from
Conifidentia~l, 8/23/54)

An investigation has been made in the Langley high-
spee~d 7- by 10l-foot tunnel to determine the aero-
dynamc characteristics at high subsodc speeds of a
modified delta wing with an aspect ratio of 3, a taper
ratio of 0.313, a sWeeplack of 27.70, and an NACA
641A012 airfoil section in combinationl with a fuse-
lage. Lift, drag, and pitching-moment characteris-
ties with free- and fixed-transition are presented
through a Mach number range of 0.40 to 0.00.


NACA RM L50C22

AERODYNAMIC AND LAERAL-CONTROL CHAR-
ACTERISTICS OF A 1/28-SCALE MODEL OF THE
BELL X-1 AIRPLANE W~ING-FUSELAGE COMBINA-
TION. TRANSONIC-BUMP METHOD. Vernard E.
Lockwrood. May 5,1950. 28p. diagrs., tab. (NACA
RM L50C22) (Declassified from Confidential,
8/23/54)

This paper presents the lateral-control characterts-
tics and the pitching-moment characteristics of a
1/28-scale model of the X-1i wing-fuselage configura-
tion. The tests were made in the transonic speed
range from a Mach nuber of 0.60 to 1.15 in the
Langle~y high-speed 7- by 10-foot tunnel utilizing the
transonic bump. Comparisons have been made be-
tween the wind-tunl results and the flight-test re-
sults of the rolling effectiveness, pitching-moment







NACA
RESEARCH ABSTRACTS NO. 70



NACA RM L50D20

FLIGHT INVESTIGATION OF THE AILERON CHAR-
ACTERISTICS OF THE DOUGLAS D-558-I AIR-
PLANE (BUAERO NO. 37972) AT MACH NUMBERS
BETWEEN 0.6 AND 0.89. Jim Rogers Thompson,
William S. Roden and John M. Eggleston. May 26,
1950. 23p. diagrs., photos., tab. (NACA
RM L50D20) (Declassified from Confidential,
8/23/54)

Measurements of the aileron effectiveness of the
Douglas D-558-1 airplane (BuAero No. 37972) indi-
cate that there is no change in aileron effectiveness
with Mdach number between Mach nubers of 0.6 and
0.89.




NACA RM L50D21

INVESTIGATION OF THE NACA 3-(3)(05)-05
EIGHT-BLADE DUAL-ROTIATING PROPELLER AT
FORWARD MACH NUMBERS TO 0.925. Robert J.
Platt, Jr. and Robert A. Shumaker. June 19, 1950.
42p. diagrs., photo. (NACA RM L50D21) (Declas-
sified from Confidential, 8/23/54)

Force-test results are presented for the NACA 3-
(3)(05)-05 eight-blade dual propeller, which was
designed for a blade angle of approximately 750. The
tests covered a blade-angle range from 550 to 80o at
Mach numbers to 0.925. Some data on the effect of
small changes in the rear-propeller blade angle are
included. Good efficiencies were obtained at high
subsonic Mach numbers by operation at high blade
angles.



NACA RM L50D24

A TECHNIQUE UTILIZING ROCKCET-PROPELLED
TEST VEHICLES FOR THE MEASUREMENT OF THEI
DAMPING IN ROLL OF STING-MOUNTED MODELS
AND SOME INITIAL RESULTS FOR DELTA AND UN-
SWEPT TAPERED WINGS. William M. Bland, .Tr.
and Carl A. Sandahl. Jme 13, 1950. 31p. diagrs.,
photos., tab. (NACA RM L50D24) (Declassified
from. Confidential, 8/23/54)

A technique with which the damping In roll of sting-
mounted models can be measured over the subsonic,
transonic, and supersonic speed range with rocket-
propelled test vehicles is described. Initial results
for delta and unswept tapered wings are presented
and compared wpith theory.



NACA RM L50D28

EFFECT OF COMPRESSIBILITY AND CAMBER AS
DETERMINED FROM AN INVSTIGATION OF THE
NACA 4-(3)(08)-03 AND 4-(5)(08)-03 TWO-BLADE
PROPELLERS UP TO FORWARD MACH NUMBERS
OF 0.925. Melvin M. Carmel, Francis G. Morgan,
Jr. and Domenic A. Coppolino. June 29, 1950. 92p.
diagrs., photo. (NACA RM LSOD28) (Declassified
from Confidential, 8/23/54)


characteristics, and lateral centers of pressure. The
estimations of rolling effectiveness were made using
the damping in roll and aileron effectiveness experi-
mentally determined.



NACA RM L50C28

MAXIUM-LIFT INVESTIGATION OF A 1/40-SCALE
X-1 AIRPLANE WING AT MACHI NUMBERS FROM
0.60 TO 1.15. Thomas R. Turner. April 21, 1950.
18p. diagrs. (NACA RM L50C28) (Declassified from
Confidential, 8/23/54)

This paper contains the lift, drag, and pitching-
moment characteristics up to a maximum lift of a
1/40O-scale model of the X-1 airplane wing (8-
percent-thick wing), also some comparisons with
flight results. The Mach number varied from 0.60
to 1.15 with the Reynolds number varying from
415,000 to 533,000.



NACA RM L50D05

AERODYNAMIC CHARACTERISTICS AT A MACH
NUMBER OF 1.25 OF A 6-PERCENT-THICK TRIAN-
GULAR WING AND 6- AND 9)-PERCENT-THICKC TRI-
ANGULAR WINGS IN COMBINATION WIT A FUSE-
LAGE. WING ASPECT RATIO 2.31, BICONVEX AIR-
FORL SECTIONS. Albert W. Hall and Garland J.
Morris. May 5, 1950. 22p. diagrs., photo., 2 tabs.
(NACA RM L50DO5) (Declassified from Confidential,
8/23/54)

Tests were made by the NACA wing-flow method on
two wing-fuselage models with delta wings of aspect
ratio 2.31, 6- and 9-percent-thick biconvex sections,
and on the 6-percent-thick wing alone. Lift, drag,
pitching-moment, and angle-of-attack measurements
are presented for these configurations and some com-
parisons are made with subsonic and supersonic data
from other sources. The test Mach number was 1.25
and the Reynolds nuber was about 8.8 x 105*




NACA RM L50D19

LOW-SPEED INVESTIGATION OF DEFLECTABLE
WING-TIP ELEVATORS ON A LOW-ASPECT-RATIO
UNTAPERED 450 SWEPTBACKC SEMISPAN WING
WITH AND WITHOUT AN END PLATE. Jack
Fischel and William~ M. O'Hare. June 1, 1950. 21p.
diagrs., photo. (NACA RM L50D19) (Declassified
from Confidential, 8/23/54)

Results and discussion are presented of a low-speed
investigation of triangular- and parallelogramzmic-
plan-form deflectable wing-tip elevators on an un-
tapered 450 sweptback semispan wing with and wcrith-
out a rectangular end plate (simulating a vertical fin)
mounted on the wing inboard of the elevators. Lift,
drag, and pitching-moment data were obtained
through a large angle-of-attack range and a large
elevator-deflection range for each elevator plan form
on the plain wing and the wing with end plate. Esti.
mated pitching-moment data were compared with
experimental data for the two elevator plan forms.







NACA
RESEARCH ABSTRACTS NO. 70

Section-thrust-coefficient data are presented for
forward Mach nungbers up to 0.925 for the NACA
4-(3)(08)-03 propeller for blade angles of 550 to 650
and for thet NACA 4-(5)(08)-03 propeller for blade
angles of 400 to 650. A comparison of these results
with force-test results is made.



NACA RM L50E09

LAT'ERAL-CONTROL INVESTIGATION OF FLAP-
TYPE AND SPOILER-TYPE CONTROLS ON A WING
WITH QUARTER-CHORD-LINE SWEEPBACK OF
600, ASPECT RATIO 2, TAPER RATIO 0.6, AND
NACA 65A006 AIRFOIL SECTION. TRANSONIC-
BUMP METHOD. Alexander D. Hammond. July18,
1950. 26p. diagrs. (NACA RM L0ED9) (Declas-
sified from Confidential, 8/23/54)

This paper presents the results of a lateral-control
investigation of a 30-percent-chord flap-type control
and a plain spoiler-type control of 5 -percent-chord
projection, each having various span and ;spanwise
locations on a wing with a quarter-chord line swept
back 600, aspect ratio 2, taper ratio 0.6, and an
NACA 65A000 airfoil section parallel to free stream.
Rolling: and pitching moments and lift of the semispan
wing-fuselage were obtained through a Mach number
range of 0.17 to 1.15 with flap-type controls, and
lift, drag, pitching moment, rolling moment, and
yawing moment of the semispan wing-fuselage were
obtained through a Mach number range of 0.6 to 1.15
with the spoiler-type controls.


NACA RM L50E19a

FLIGHT INVESTIGATION AT MACH NUMBERS
FROM 0.6 to 1.7 TO DETERMINE DRAG AND BASE
PRESSURES ON A BLUNT-TRAILING-EDGE AIR-
FOIL AND DRAG OF DIAMOND AND CIRCULA-
ARC AIRFOILS AT ZERO LIFT. John D. Morrow
and Elis Katz. August 11, 1950. 25p. diagrs.,
photos. (NACA RM L50E10a) (Declassified from
Confidential, 8/23/54)

Results of an exploratory free-flight investigation at
zero lift of several rocket-powered drag-research
models having rectangular 6-percent-thick wings are
presented for a Mach number range of 0.6 to 1.7.
Wings having diamond, circular-arc, and blunt-
trailing-edge airfoil sections were tested. Pres-
sures over the base of the blunt-trailing-edge airfoil
were measured. The drag of all the models were
measured and are compared with theory in this
paper,


NACA RM L50E26

DAMPING IN ROLL OF RECTANGULAR WINGdS OF
SEVERAL ASPECT RATIOS AND NACA 65A-SERIES
AIRFOIL SECTIONS O1F SEVERAL THICKNESS
RATIOS AT TRANSONIC AND SUPERSONIC SPEEDS
AS DETERMINED WITH ROCKET-POWERED
MODELS. James L. Edmondson. August 24, 1950.
16p. diagrs. (NACA RM L50E26) (Declassified
from Confidential, 8/23/54)

The damaping in roll has been determined for a series
of rectangular wings of various aspect ratios and
thickness ratios through use of rocket-powered


27


models and the canted-nozzle technique. An empir-
ical thickness correction factor has been derived to
be used with existing linear supersonic damping-in-
roll theory to allow closer prediction of experimental
damping for limited cases,





NACA RM L50F01

RESULTS OF FLIGHT TESTS TO DETERMINE THE
ZERO-LIFT DRAG CHARACTERISTICS OF A 600
DELTA WING WITH: NACA 65-006 AIRFOIL SEC-
TION AND VARIOUS DOUBLE-WEDGE SECTIONS
AT MACH NUMBERS FROM 0.7 TO 1.6. Clement J.
Welsh. August11, 1950. 15p. diagrs., photo.
(NACA RMi L50F01) (Declassified fromn Confidential,
8/23/54)

Results of an exploratory free-flight investigation at
zero lift of several rocket-powered drag research
models equipped with 600 sweptback delta wings are
presented for a Mach number range from about 0.70
to 1.60. The airfoil sections tested included the
NACA 65-006 and a series of double-wedge sections
with various positions of maximum thickness.





NACA RM L50Fl9

LOW-SPEED TESTS OF A MODEL SIMULATING
THE PHENOMENON OF CONTROL-SURFACE
BUZZ. William H. Phillips and James J. Adams.
August 16, 1950. 16p. diagrs., photo. (NACA
RM L50F19) (Declassified from Confidential,
8/23/54)

Low-speed tests have been made of an airfoil model
with a freely hinged flap connected to spoilers which
passed through slots in the airfoil ahead of the hige
line. Under certain conditions, continuous oscilla-
tions of the flap, similar to buzz, were obtained.
This result indicates that flow separation may be an
important factor in producing buzz at transonic
speeds.



NACA RM L50F19a

SPIN AND) RECOVERY CHARACTERISTICS OF A
MODEL OF A FIGHTER TYPE OF AIRPLANE WITH-
OUT A HORIZONTAL TAIL, AND HAVING EITHER A
SINGLE VERTICAL TAIL OR TWIN VERTICAL
TAILS. Lawrence J. Gale and Norman E.
Pumphrey. July 25, 1950. 23p. diagram photos.,
2 tabs. (NACA RM L50Fl9a) (Declassified from
Confidential, 8/23/54)

The investigation indicated similar spin and recovery
characteristics for either tail configuration tested at
a given mass distribution. For a mass distribution
chiefly along the wings, the vertical tail surfaces
were not adequate for recovery from the spin. When
the mass was distributed chiefly along the fuselage,
however, either vertical-tail configuration, when in
a rearwtpard position, was effective in satisfactorily
terminating the spin.







28


NACA RM L50F26
National Advisory Committee for Aeronautics.
MEASUREMENTS OF THE EFFECT OF TRAILING-
EDGE THICKNESS ON THE ZERO-LIFT DRAG OF
THIN LOW-ASPECT-RATIO WINGS. John D.
Morrow. August 14, 1950. 12p. diagrs., photo.
(NACA RM L50F26) (Declassified from Confidential,
8/23/54)

Results of an exploratory free-flight investigation at
zero lift of several rocket-powered drag-research
models having 4-percent-thick wings of 0.423 taper
ratio are presented for a Mach number range of 0.7
to 1.6. Four wings having trailing edges of different
thickness were tested. The drag of all the models
was measured and is compared with calculated
values in this paper,


NACA RM L50F26a

SOME CALCULATIONS OF THE LATERAL RE-
SPONSE OF TWO AIRPLANES TO ATMOSPHERIC
TUBLENCE WITH R LTIONiTO THEl LTRAL

1950. 24p. diagrs., 2 tabs. (NACA RM LOF26a)
(Declassified from Confidential, 8/23/54)

Calculations are made of the lateral response to
representative time histories of atmospheric turbul-
ence for two airplanes having widely different dy-
namic properties; explanations for their difference
in behavior are given. The results are discussed in
relation to lateral snaking.



NACA RM L50F30

A THEORETICAL INVESTIGATION OF THE INFLU-
ENCE OF AUXILIARY DAMPING IN PITCH ON THE
DYNAMIC CHARACTERISTICS OF A PROPORTION-
ALLY CONTROLLED SUPERSONIC CANARD MIS-
SILE CONFIGURATION. WHalter C. Nelson and
Anthony L. Passera. August 25, 1950. 46p.
diagrs., photo., 3 tabs. (NACA RM L50F30)
(Declassified from Confidential, 8/23/54)

A theoretical investigation is presented of the dynam-
ic characteristics of a supersonic canard missile
configuration with auxiliary damping in pitch obtained
from rate gyro-servo control. Variations in Mach
number, altitude, and static margin are considered.
Satisfactory system performance is obtained
throughout a Mach number and altitude range when
additional damping is included in the missile.



NACA RM L50F30a

FLIGHT TESTS AT SUPERSONIC SPEEDS TO
DETERMINE THE EFFECT OF TAPER ON THE
ZERO-LIFT DRAG OF SWEPTBACK LOW-ASPECT-
RATIO WINGS. Murray Pittel. September 5, 1950.
23p. diagrs., photos. (NACA RM L50F30a) (Declas
sified from Confidential, 8/23/54)

Comparison of experimental flight-test results with
theoretical calculations to determine the effect of
wing taper and aspect ratio on the zero-lift drag of
wings with thin double-wedge sections at supersonic
speeds is presented.


NACA
RESEARCH ABSTRACTS NO. 70

NACA RM L50G03

INVESTIGATION OF FLAP-TYPE AILERONS ON AN
UNTAPERED WING HAVING AN ASPECT RATIO OF
3.7, 450 SWEEPBACK, AND AN NACA 65A\009 AIR-
FOIL SECTION. TRANSONIC-BUMP METHOD.
Richard G. MacLeod. August 23, 1950. 18p.
diagrs. (NACA RM L50GO3) (Declassified from.
Confidential, 8/23/54)

This paper presents the results of an investigation to
determine the lateral control characteristics of 20-
percent-chord flap-type control surfaces of various
spans on a semispan wing-fuselage model by the
transonic-bump method. The model employed a
wing with a sweepback of 450, an aspect ratio 3.7, a
taper ratio of 1.0, and an NACA 65A009 airfoil sec-
tion parallel to the free stream. Rolling moments
were obtained through a small range of angle of at-
tack and control-surface deflections. Lift data on
the complete model are also included. The experi-
mental results were in good agreement with those
predicted from low-speed theory and other experi-
mental data atea Mac number ofe0 le and mae ra-

firl~ydconstant throughout the Mach number range





NACA RM L50G11

LOW-SPEED STATIC STABILITY CHIARACTEI-
TICS OF A CANARD MODEL WITH A 450 SWEPT-
BACK WING AND A 600 TRIANGULAR HORIZONTAL
CONTROL SURFACE. John W. Draper.
September 6, 1950. 43p. diagrs., photo., 2 tabs.
(NACA RM L50G11) (Dleclassified from Confidential,
8/23/54)

This paper contains results of an investigation made
to determine the low-speed static stability character-
istics of a canard model having a 450 sweptback un-
tapered wing and a 600 triangular-plan-formn hori-
zontal control surface.





NACA RM L50G13a

INVESTIGATION AT TRANSONIC SPEEDS OF A
35-PERCENTI-CHORD AILERON ON A TAPERED
WEDGE-TYPE WING OF ASPECT RATIO 2.5 WITH
AND WITHOUT A FUSELAGE. Thomas R. Turner
and Joseph E. Fikes. September, 1950. 25p.
diagrs. (NACA RM L50G13a) (Declassified from
Confidential, 8/23/54)
This paper presents the results of an investigation to
determine the effects of fuselage diameter on some of
the aerodynamic characteristics of an unswept wing
having a modified double-wedge section, an aspect
ratio of 2.5, and a taper ratio of 0.625. Rolling mo-
inent coefficients for a 35-percent-chord aileron of
various spans along with some lift, drag, and pitching
moment coefficient data are presented for the wing
with a fuselage of two different diameters and without
a fuselage through a Mach number range from 0.60 to
1.18. Estim~ated and expe~rimnental values of aileron
effectiveness are presented at a Mach numberof0.60.








NACA
RESEARCH ABSTRACTS NO. 70

NACA RMI L50G14

AE~RODIYNAMIC CHARACTERISTICS AT A MLACH
NUMBER OF 1.38 OF FOUR WINGS OF ASPECT
RATIO 4 HAVING QUARTER-CHORD SWEEP
ANLES OF 00, 350, 45o, AND 600. William B.
Kemp, Jr., Kenneth W. Goodson and Robert A. Booth,
October 10, 1950. 41p. diagrs., photos., tab.
(NACA RM L50G14) (Declassified from Confidential,
8; 23 54)

This paper presents a description of the Langley 6-
inch supersonic tunnel together with the results of
an investigation made at a Mach nube of 1.38 of a
series of wings having an aspect ratio 4, taper ratio
0.6, and an NACA 65A006 airfoil section. Lift, drag,
pitching-moment, and bending-moment data were ob-
tained for wing alone and wing-bodfy configurations.
Comparisons are made between experimental and
theoretical lift-curve slopes and stability character-
istics.

NACA RM L50G14a

THE EXTERNAL-SHOCK DRAG OF SUPERSONIC
ILoT .KAu G SUBSONIC 2ENT ACE8 LdO r.,
photos. (NACA RM L50G14a) (Declassified from
Confidential, 8/23/54)

The external-shock drag at a Mach number of 2.70
has been determined by two methods for a circular
inlet having low external compression at various
mass-flow conditions and for a circular inlet having
high external compression. The results indicate a
large increase in external shock drag for both inlets
at subdesign mass flows, with the drag of the external
compression type being considerably higher for com-
parable mass-flow reduction than that of the inlet
having low external compression. A simple approxi-
mate method of calculating the external-shock drag is
d scribed wic aise re esil dplield than t pro-
flow and the shock configuration are both known.


NACA RM L50G18

THE CALCULATION OF THE PATH OF A JETTI.
SON~ABLE NOSE SECTION. Roscoe H. Goodwrin.
September 7, 1950. 35p. diagrs. (NACA RM L50G18)
(Declassified from Confidential, 8/23/ 54)

A method is presented for calculating the path of a
jettisonable nose section at any speed by means of
successive approximations (solved graphically) using
static aerodynamic characteristics. Comparisons
with exqperimnentally determined paths are presented,
and the method is applied to the problem of finding
the path at a nose section that is jettisoned by use of
rockets.


NACA RM L50G24

FLIGHT MEASUREMENTS OF DRAG AND BASE
PRESSURE OF A FIN-STABILIZED PARABOLIC
BODY OF REVOLUTION (NACA RM-10) AT DIF-
FERENT REYNOLD NUMBERS AND AT MACH
NUMBERS FROM 0.0 TO 3.3. H. Herbert Jackson,
Charles B. Rusey and Leo T. Chauvin.
September 1, 190. 21p. diagrs., photos. (NACA
RM L50G24) (Declassifld from Conildental.
8/23 54)


2p


Free-flight. tests at supersonic speeds have been
made to determine the Reynolds nuber effects on
total drag and base drag of a fin-stabilized parabolic-
arc body of revolution having a body fineness ratio of
12.2 and designated the NACA RMn-10 configuration.
The Reynolds number range of 14 x 106 to 210 x 106
was obtained by testing full-scale and half-scale
models through the Mach number range from 0.9 to
3.3.


NACA RMb L50HO03

PROPELLER SECTION AERODYNAMIC CHARAC-
TERISTICS AS DETERMIED BY MEASURING THIE
SECTION SURFACE PRESSURES ON AN NACA
10-(3)(08)-03 PROPELLER UNDER OPERATING
CONDITIONS. Albert J. Evans. November 8, 1950.
162p. diagrs., 10 tabs (NACA RM L50HO3) (De-
classified from Confidential, 8, 23 54)

This investigation presents propeller section aero-
dynamic characteristics obtained by measurement of
the section surface pressure distribution at nine
radial stations on an operating propeller blade. The

toenesioa 1 roil and te orman e of M h
numbers from 0.20 to 1.15 for angles of attack from
-lo to 120 for NACA 16-series airfoils.


NACA RM L50HOT

LONGITUDINAL STABILITY AND CONTROL CHAR-
ACTEIS~TICS AT HIGH-SUBSONIC SPEEDS OF TWO
MODELS OF A TRANSONIC RESEARCH AIRPLAE
WIITH WINGS AND HORIZONTAL TAILS OF ASPECT
RATIOS 4.2 AN 2. Arvo A. Luoma and John B.
Wright. September 29, 1950. 134p. diagrs., photos.,
3 tabs. (NACA RM L50H07) (Declassified from
Confidential, 8, 23 54)

This paper contains longitudinal stability and control
characteristics of two 1/16i-scale models of a tran-
sonic research airplane tested with no nose-inlet
flow. Aspect ratio was the main variable. Various
horizontal-tail incidences and elevator deflections
were tested. The tests were made up to a Mach num
ber of approximately 0.95 in the Langley 8-foot high-
speed tunnel.


NACA RM L50BO9

MEASUREMENTS OF AERODYAMC CHARACTER-
ISTICS OF A 350 SWEPTBACKI NACA 65-009 AI-
FOIL MODEL WITH 1/4-CHORD FLAP HAVING A
31-PERCENT-FLAP-CHORD OVERH~ANG BALANCE
BY THE NACA WING-FLOW METHOD. H~arold I.
Johnson and Harold R. Goodman. September 25,
1950. 38p. diagrs., photo. (NACA RM L50H09)
(D~eclassified from Confidential, 8/23 54)

Lift, pitching-moment, and hinge-mnoment data ob-
tained from wing-flow tests of a low-aspect-ratio
sweptback-airfoil model havig a full-span 1/4l-chord
overhn-balanced flap are presented. Mach number
range was 0.55 to 1.15, Comparisons are made with
previo~usly unpublished data from tests of a plain-flap
model. Conclusion is that 31-percent-flap-chord
overhang balance tested was relatively ineffective for
reducing hinge moments below a Mach number of 1.00
and totally ineffective between Mach nubers of 1.00
and 1.15.






30


NACA RM L50H11

EFFECTS OF SWEEP ON THE MAXIMUM-LIFT
CHARACTERISTICS OF FOUR ASPECT-RALTIO-4
WINGS AT TRANSONIC SPEEDS. Thomas R. Turner.
October 3,1950. 25ip.diagrs. (NACA RM L50H11)
(Declassified from Confidential, 8/23/54)

An investigation at transonic speeds has been made
to determine the effect of wing sweep on the maxi-
mum lift characteristics of a series of wings having
an aspect ratio of 4, a taper ratio of 0.6, and the
quarter-chord line swept back Go, 350, 450, and 600.
The Mach number varied from 0.61 to 1.20 with a
Reynolds number variation from 380,000 to 460,000.
Lift data are presented from approximately zero lift
to beyond maximum lift. Drag and pitching-moment
data are also presented.


NACA RM L50H15

LOW-SPEED INVESTIGATION OF A SEMISUB-
MERGED AIR SCOOP WITH AND WITHOUT
BOUNDARY-LAYER SUCTION. P. Kenneth Pierpont
and Robert R. Howell. February 23, 1951. 46p.
diagrs., photos., 2 tabs. (NACA RM L50HI15) (De-
classified from Confidential, 8/23/54)

This investigation presents results at low-speed
tests of a half-submerged inlet consisting of a con-
ventional scoop located in a depression in the fuse-
lage surface. Boundary-layer-control systems in-
vestigated are shown to provide increases in impact-
pressure ratio up to 8 percent at Vi/Vo = 0.6. The
impact-pressure recovery of a partly submerged inlet
appears to exceed that of a submerged inlet in the
inlet-velocity-ratio range suitable for high-speed
operation.


NACA RM L50H22

TRANSONIC DRAG CHARACTERISTICS OF A WING-
BODY COMBINATION USING A THIN TAPERED
WING OF 450 SWEEPBACK. Max C. Kurbjun and
Stanley Faber. September 28,1950. 14p. diagrs.,
photo., tab. (NACA RM L50H22) (Declassified from
Confidential, 8/23/54)

Contains drag measurements by the free-fall method
for a wing-body combination and its components em-
ploying a wing of 0.2 taper ratio, NACA 65003 airfoil
section (perpendicular to midchord line), 450 sweep-
back (midchord line), and a fineness-ratio-12 body.


NACA RM L50H23

PRELIMINARY EMPIRICAL DESIGN REQUIRE-
MENTS FOR THE PREVENTION OF TUMBLING OF
AIRPLANES HAVING NO HORIZONTAL TAILS.
Robert L. Bryant. October 11, 1950. 23p. diagrs.,
2 tabs. (NACA RM L50H23) (Declassified from
Confidential, 8/23/54)

An investigation has been made of the design charac-
teristics and loadings that are conducive to the
tumbling of airplanes that have no horizontal tails.
Preliminary empirical design requirements based on
model tests of 18 different configurations are pre-
sented. A brief explanation of the phenomenon of
tumbling is appended.


NACA
RESEARCH ABSTRACTS NO. 70

NACA RM L50H24

AN INVESTIGATION OF THREE TRANSONIC FUSE-
LAGE AIR INLETS AT MACH NUMBERS FROM 0.4
TO 0.94 AND AT A MACH NUMBER OF 1.19, Robert
E. Pendley, Harold L. Robinson and Claude V.
Williamns. November 7, 1950. 51p. diagrs., photos.,
3 tabs. (NACA RM L50H24) (Declassified from
Confidential, 8/23/54)

Measurements of internal-flow pressure recovery,
external-surface pressures, and external drag are
presented for three fuselage-side inlets designed for
use at transonic speeds. The test Mach number and
inlet-velocity ratio ranged from 0.4 to 1.19 and from
0 to 1.9, respectively. The investigation showed that
the maximum value of the impact-pressure recovery
was high at all Mach numbers. The use of a higher
critical Mach number external shape increased the
supercritical drag-rise Mach number, decreased the
supercritical drag, and reduced the drag at the super-
sonic Mach number.


NACA RM L50HI28a

EFFECT OF AN END PLATE ON THE AERODY-
NAMIC CHARACTERISTICS OF A 20.550 SWEPT-
BACK WING W~ITH AN ASPECT RATIO OF 2.67 AND
A TAPER RATIO OF 0.5. TRANSONIC-BUMP
METHOD. James M. Watson. December 21, 1950.
15p. diagrs., photo. (NACA RM L50HI28a) (De-
classified from Confidential, 8/23/54)

This paper contains the results of an investigation of
two wings having an aspect ratio of 2.67, a sweep-
back of 20.550, and a taper ratio of 0.5, one wing
with and one wing without an end plate. A Mach
number range of 0.60 to 1.18 was attained by use of
the transonic-bump technique.



NACA RM L50H30a

DAMPING IN YAW AND STATIC DIRECTIONAL,
STABILITY OF A CANAR AIRPLANE MODEL AND
OF SEVERAL MODELS HAVING FUSELAGES OF
RELATIVELY FLAT CROSS SECTION. Joseph L.
Johnson. October 16, 1950. 20p. diagrs., tab.
(NACA RM L50H30a) (Declassified from. Confiden-
tial, 8/23/54)

Contains results of damping-in-yaw and static-
directional-stability tests for flat-tuselage models
having major axis horizontal and vertical, for a flat-
fuselage model with major axis horizontal in combi-
mation with a 450 sweptback wing, and for a canard
model having a triangular horizontal control surface
and a 450 sweptback wing. The effect of a vertical
tail located at the rear of the fuselage was deter-
mined for each model investigated.



NACA RM L50I01

LOW-SPEED WIND-TUNNEL INVESTIGATION OF A
TRIANGULAR SWEPTBACK AIR INLET IN THE
ROOT OF A 450 SWEPTBACK WING. Arvid L.
Keith, Jr. and Jack Schiff. November 6, 1950. 71p.
diagrs., photos., 5 tabs. (NACA RM L50I01) (De-
classified from Confidential, 8/23/54)








NACA
RESEARCH ABSTRACTS NO. 70


Results of a low-speed study of a 450 sweptback wing-
root air-inlet configuration believed suitable for tran-
sonic speed airplanes are presented. The inlet con-
figuration lift and drag characteristics are compared
with those of a basic model. Boundary-layer growth
along the fuselage nose, inlet total-pressure recov-
eries, and static-pressure distributions over the in-
let and wing surfaces are presented for wide ranges
of inlet-velocity ratio and angle of attack.


NACA RM L50IOS

EFFECTS ON THE LATERAL OSCILLATION OF
FIXING THE RUDDER AND REFLEXING THE
FLAPS ON THIE BELL X-1. AIRPLANE. Hubert M.
Drake. December 11, 1950. 14p. diagrs., photo.
(NACA RM L50I05) (Declassified from Confidential
8/23/54) '

In flight tests of the Bell X-1 airplane at a Mach
number of 0.85 it has been found that fixing the
rudder reduced the amplitude of the snaking oscilla-
tion, but did not eliminate it. It was also found that
reHexn the win flaps i cang ahe dhmliaton of
lateral stability, but had only a small effect on the
snakig oscillation.



NACA RMI L50IO8a

FLIGHT INVESTIGATION AT MAC NUMBERS
FROM 0.8 TO 1.5 TO DETERMINE THE EFFECTS
OF NOSE BLUNTNESS ON THE TOTAL DRAG OF
TWO FIN-STABILIZED BODIES OF REVOLUTION.
RogerG. Hart. October 16, 1950. 12p. diagrs.,
photos., tabs. (NACA RM L50108a) (Declassified
from Confidential, 8 23 '54)

Values of total-drag coefficient were measured for
two fin-stabilized, blunt-nose bodies of revolution in
free flight at Mach numbers from 0.8 to 1.5. The
blunt-nose bodies were designed by rounding off the
sharp noses of bodies, having nose fineness ratios of
about 3-1/2, to radit equal to about 1/4 the maxi-
mum radii. No increase in the drag of either body
was found'



NACA RM L50I28a

FLIGHT MEASUREMENTS OF BASE PRESSURE ON
BODIES OF REVOLUTION WITH AND WITHOUT
SIMULATED ROCKET CHAMBERS. Robert F. Peck.
November 16, 1950. 21p. diagrs., photos. (NACA
RM L50128a) (Declassified from Confidential,
8 23 54)

Base pressures were measured on fin-stabilized
bodies of revolution with and without rocket chambers
and with and without a converging afterbody. At
Mach nubers between 0.7 and 1.2, the results show
that the presence of a "cold" rocket chamber in-
creased the pressure (less suction) over the center
portion of the bases. The effects of rocket chambers
on pressures near the edge of the bases waere not as
consistent throughout the Mach number range nor as
appreciable at most speeds as were the effects on
pressures measured on the center line.


NACA RM LSOJO2

A PRELIMINARY FLIGHT INVESTIGATION OF THE
EFFECTS OF VORTEX GENERATORS ON SEPARA-
TION DUE TO SHOCK. Lindsay J. Lina and Wilmer
H. Reed, 2I. November 30, 1950. 30p. diagrs.,
photos., tab. (NACA RM L50JO2) (Declassified
from Confidential, 8 23 '54)

A preliminary flight investigation showed that several
configurations of vortex generators mounted on the
upper surface of a modified wing of an F-51D air-
plane reduced separation caused by shock.


NACA RM L50J05

A WIND-TUNNL INESTIGATION OF THE AERO-
DYNAMIC CHARACTERISTICS OF A FULL-SCALE
SWEPTBACK PROPELLER AND TWO RELATED
STRAIGHT PROPELLERS. Albert J. Evans and
George Liner. January 4, 1951. 102p. diagrs.,
photos., tab. (NACA RM L50J05) (Declassified
from Confidential, 8/31/54)

Prpele a 2 ynmic b arac ei tic are pr nted

-057-27 two-blade swept propeller and two related
straight propellers is included. An increase of about
6 percent in the value of helical tip Mach number at
which compressibility losses become manifest was
realized by the use of sweep. Analysis of the results
wiarns against pseudo gains attained by the use of
sweep in propellers at the expense of overall per-
formance.


NACA RM L50J06

INVESTIGATION OF A 42.7o SWEPTBACK WING
MODEL TO DETERMINE THE EFFECTS OF
TRAILING-EDGE, THICKNESS ON THE AILERON
HINGE-MOMENT AND FLUTTER CHARACTER-
ISTICS AT TRANSONIC SPEEDS, Robert F.
Thompson. December 26, 1950. 42p. diagrs.,
photos., 2 tabs. (NACA RM L50JO6) (Declassified
from Confidential, 8/31 5411

A wind-tunnel investigation of a semnispan wing-
fuselage model having 42. 70 of sweepback of the wing
leading edge was made through a speed range to a
Mach number of 0.60 to 1.175. The 0.20-chord out-
board aileron was tested with three trailing-edge to
hinge-line thickness ratios (t = 0, t = 0.5, and
t = 1.0) to determine the effects on hinge moments
and one-degree-of-freedom. aileron flutter. Pre-
sented are hingle-momntrn data, hinge-moment pa-
rameters, and allerorn tree-floating characteristics.
A comparison of the flutter frequency is made with
two previously published empirical analyses.


NACA RM L50J10

TABULATED PRESSURE COEFFICIENTS AND)
AERODYNAMIC CHARACTERISTICS MEASURED IN
FLIGHT ON THE WING OF THE DOUGLAS D-558-I
AIRPLANE FOR A 1 g STALL, A SPEED RUN TO A
MACH NUMBER OF 0.90, AND A WIND-UP TURN
AT A MACH NUMBER OF 0.86. Earl R. Keener and
Mary Pierce. December 15, 1950. 40p. diagrs.,
photos., 5 tabs. (NACA RM L50J10) (Declassified
from Confidential, 8/31/54)






NACA
RESEARCH ABSTRACTS NO. 70


NACA RM L50J19

EFFECTS OF SWEEP ON THE DAMPING-IN-ROLL
CHARACTERISTICS OF THREE SWEPTBACK
WINGS HAVING AN ASPECT RATIO OF 4 AT
TRANSONIC SPEEDS. Vernard E. Lockwood.
December 14, 1950. 23p. diagrs. (NACA
RM L50J19) (Declassified from Confidential,
8/31/54)

The damping-in-roll characteristics of three wings
of aspect ratio 4 and taper ratio 0.6 with sweep
angles of Do, 350, and 450 at the quarter-chord line
and an NACA 65A006 airfoil section have been
determined through the Mach number range from 0.6
to 1.15 and an angle-of-attack range from 00 to
approximately 70 in the Langley 7- by 10-foot-tunnel
transonic bump by the twisted-wing method.
Comparisons are included with the results obtained
at subsonic Mach numbers by the free-roll method on
a series of similar wings.


NACA RM L50J20

LOW-SPEED INVESTIGATION OF THE EFFECT OF
SEVERAL FLAP AND SPOILER AILERONS ON THE
LATERAL CHARACTERISTICS OF A 47.50
SWEPTBACK-WING FUSELAGE COMBINATION
AT A REYNOLDS NUMBER OF 4.4 x 106. Jerome
Pasamanick and Thomas B. Sellers. December 8,
1950. 57p. diagrs., photo. (NACA RM L50J20)
(Declassified from Confidential, 8/31/54)

Presents results of an investigation in the Langley
full-seale tunel of various plain spoiler and flap-
type ailerons on a wing-fuselage combination. The
wing leading-edge sweep was 47. 50, the aspect ratio
was 3.4, the taper ratio was 0.51, and the airfoil
sections were NACA 641A112. The data include the
effects of aileron span, location, and trailing-edge
thickness and of spoiler span, location, and projec-
tion on the longitudinal and lateral characteristics at
zero yaw for a range of angle of attack at a Reynolds
number of 4.4 x 106,


NACA RM L50J30

FLIGHT INVESTIGATION OF THE EFFECT OF
SIDESLIP ON THE PRESSURE AT THE STATIC
ORIFICES OF THE BOEING B-29 AIRPLANE. Robert
G. Chilton and B. Porter Brown. April 11, 1951. 12p.
diagrs. (NACA RM L50J30) (Declassified from
Confidential, 8/31/54)

The effect of sideslip on the measurement of static
pressure obtained from orifices located on each side
of the fuselage near the nose of an airplane is shown
by measurements taken in steady and oscillatory
sideslips. The result is discussed with respect to
its effect on the indications of change and rate of
change of altitude of an airplane on a bombing run.



NACA RM L50K01a

EFFECTS ON THE SNAKIG OSCILLATION OF THE
BELL X-1 AIRPLANE OF A TRAIING-EDGE BULB
ON THE RUDDER. Hubert M. Drake and Harry P.
Clagett. January 16, 1951. 14p. diagrs., photo.
(NACA RM L50K01a) (Declassified from Confidential,
8/31/54)


32



Presents tabulated pressure coefficients and aero-
dynamic characteristics measured in flight on the
right wing of the Douglas D-558-I airplane for a 1 g
stall, a speed run to a Mach number of 0.90, and a
wind-up turn at a Mach number of 0.86.




NACA RM L50J12

DRAG INVESTIGATION OF SOME FIN CONFIGURA-
TIONS FOR BOOSTER ROCKETS AT MACH NUM-
BERS BETWEEN 0.5 AND 1.4. John C. McFall, Jr.
November 21, 1950. 17p. diagrs., photos., tab.
(NACA RM L50J12) (Declassified from Confidential,
8/31/54)

This paper presents the results of an investigation
using rocket-propelled free-flight models to furnish
drag data for booster drag estimates and to deter-
mine the drag of various booster fin configurations.
Model booster fins of a type extensively used by the
NACA were flown through a Mach number range of
0.5 to 1.4. Two booster plan forms were investi-
gated: one with aspect ratio 2.04, taper ratio 0.37,
and having various tie-rod bracing and the other, a
cantilever design, with aspect ratio 3.20 and taper
ratio 0.32.





NACA RM L50J17

SKIN-TEMPERATURE TELEMETER FOR
DETERMINING BOUNDARY-LAYER HEAT-
TRANSFER COEFFICIENTS. Clifford L. Fricke
and Francis B. Smith. March 15, 1951. 22p.
diagrs. (NACA RM L50J17) (Declassified from
Confidential, 8/31/54)

A method of experimentally determining boundary-
layer heat-transfer coefficients by telemetering the
skin temperatures of supersonic rockets is given.
A platinum resistance wire temperature sensing
element developed to indicate accurately the rapidly
changing skin temperature is described and the over-
all accuracy of the instrumentation is discussed.





NACA RM L50J18

INVESTIGATION AT SUPERSONIC SPEEDS OF SOME
OF THE FACTORS AFFECTING THE FLOW OVER A
RECTANGULAR WING WITH SYMMETRICAL
CIRCULAR-ARC SECTION AND 30-PERCENT-
CHORD) TRAILING-EDGE F~LAP. K. R. Czarnecki
and James N. Mueller. January 2, 1951. 111p.
diagrs., photos. (NACA RM L50J18) (Declassified
from Confidential, 8/31/54)

The results of an investigation at supersonic speeds
(M = 1.62, 1.93, and 2.40) of some of the factors af-
fecting the flow over a rectangular wing with a sym-
metrical circular-are section and a 30-percent-
chord trailing-edge flap are presented. The factors
included are Mach number, wing thickness, effect of
fixing transition, flap-gap leakage, model asymnmetry,
and surface condition.








. NACA
RESEARCH ABSTRACTS NO. 70


It was found that a rudder trailing-edge bulb did not
appreciably affect the snaking oscillation over the
Mach number range from 0.75 to 1.0.



NACA RM L50K09

ESTIMATED DECELERATION OF AIRPLANE NOSE
SECTION JETTISONED AT VARIOUSi ALTITUDES
AND AIRSPEEDS. Stanley H. Scher. January 8
1951. 39p. diagrs. (NACA RM L50K09) (Declassi-
fled from. Confidential, 8/31/54)

An analytical Investigation has been made of the de-

rnos sa tion lfebein jtasn ted t M umb rs
ranging from. 0.85 to 3.0 and at altitudes ranging
from sea level to 120,000 feet'


NACA RM L50KE10

THE EFFECT OF MASS DISTRIBUTION ON THE
LOW-SPEED DYNAMIC LATERAL STABILITY AND
CONTROL CHARACTERISTICS OF A MODEL WITH
A 600 TRIANGULAR WING. Joseph L. Johnson.
March 9, 1951. 23p. diagrs., 2 tabs. (NACA
RM L50K10) (Declassified from Confidential,
8/3 1/54)

Results of an investigation to determine the effect of
mass distribution on the dynamic lateral stability and
control characteristics of a model with a 600 triangu-
lar wring are presented. The moments of inertia in
roll and yaw were increased to correspond to those of
a triangular-wing fighter airplane with wing tanks.
Flight tests and calculations were made for five dif-
ferent loading conditions.



NACA RM L50K20

THE USE OF SUCTION TO PREVENT SHOCK-
INDUCED SEPARATION IN A NOZZLE. James R.
Sterrett, Robert W. Dunning and Maurice J. Brevoort.
January 30, 1951. 64p. diagrs., photos., 2 tabs.
(NACA RIM L50K20) (Declassified from Confidential,
8/31/54)

An investigation ~was made of the use of suction to
prevent shock-induced flow separation in a nozzle
formed by a 71.5-percent-thick bump on the wall of a
channel. Various transverse and longitudinal suction-
slot arrangements and suction through porous sur-
faces were tested. All these devices were effective
in preventing separation, and certain suction-slot
arrangements reduced the total power loss, including
the power lost in the suction process.



NACA RM L50K27

AN EXPERIMENTAL STUDY AT MODERATE AND
]HIGH SUBSONIC SPEEDS OF THE FLOW OVER
WrIN;G WITH 300 AND 450 OF SWEEPBACK IN
CONJUNCTION WITH A FUSELAGE. Richard T.
Whitcomb. June 15, 1951. 56p. diagrs., photos.
(NACA RM L50K27) (Declassified from Confidential,
8/31/54)


33



A relatively extensive study is presented of the
pressure distributions, wake surveys, and tuft pat-
terns obtained at Mach numbers to 0.96 for tapered
wings with 300 and 450 of sweepback in conjunction
with a fuselage.



NACA RM L50K28

AN EXPERIMENTAL STUDY OF MODERATE AND
HIGH SUBSONIC SPEEDS OF THE FLOW OVER
WINGS WITH 300 AND 450 OF SWEEPFORWARD
IN CONJUNCTION WITH A FUSELAGE. Richard T.

(NhC L50K 8 (1 5cl~as4 fed fr 0 Cnietial,


A relatively extensive study is presented of the pres-
sure distributions and wake measurements obtained
at Mach numbers to 0.96 for tapered wings wfith 300
and 450 of sweeplorward, in conjunction with a
fuselage.




NACA RM L50LO1

AERODYNAMIC CHAATERISTICS AT TRANSONI C
SPEEDS OF A 600 DELTA WING EQUIPED WITH A
TRIANGULAR PLAN-FORM CONTRL HAVING A
SKEWED HINGE AXIS AND AN OVERHANG BAL-
ANCE. TRANSONIC-BUMP METHOD. H~arleth G.
Wiley. February 6, 1951, 31p. diagrs. (NACA
RM L50L01) (Declassified from Confidential,
8/31/54)

This paper presents the aerodynamic characteristics
of a 600 delta wing of aspect ratio 2.31, taper ratio 0,
and an NACA 65-006 airfoil section, which was
equipped with an aerodynamically balanced trianguar
control mounted on a skewed hinge axis. Lift, drag,
pitching moment, rolling moment, and hinge moment
were obtained at various angles of control deflection
and angles of attack through a Mach number range of
0.6 to 1.18.





NACA RML L50LO4

EFFECTS OF SEVERAL ARRANGEMENTS OF REC-
TANGULAR VORTEX GENERATORS ON THE1
STATIC-PRESSURE RISE THROUGH A SHORT 2:1
DIFFUSER. E. Floyd Valentine and Raymond B.
Carroll. February 20, 1951. 35ip. diagrs., photos.
(NACA RIM L50LO4) (Declassified from Confidential,
8/31/54)

Several arrangements of simple rectangular noncam-
bered vortex generators were investigated in a 2:1
area-ratio diffuser of length equal to the inlet diam-
eter and having an initial boundary-layer thickness of
5 percent of the inlet diameter. Some arrangements
actually reduced the diffuser static-pressure rise.
The effect of one of the better vortex-generator
arrangements was to increase the diffuser effective-
ness by 30 percent making it equal to that of a
diffuser of twice the length.







34


NACA RM L50LO4a

PRESSURE DISTRIBUTIONS OVER A RETRACTED
LEADING-EDGE SLAT ON A 400 SWEPTBACK WING
AT MACH NUMBERS UP TO 0.9. Jones F. Cahill
and Gale C. Oberndorfer, January 26, 1951. 36p,
diagrs. (NACA RM L50LO4a) (Declassified from
Confidential, 8/31/54)

Pressure distributions over a retracted leading-edge
slat on a 400 sweptback wing are presented for a
series of angles of attack up to maximnumn lift at Mach
numbers of 0.1, 0.4, 0.6, 0.7, 0.8, 0.85, and 0.9.
These data were obtained in the Langley low-
turbulence pressure tunnel at a constant Reynolds
number of approximately 3 x 106. The Reynolds
number was maintained approximately constant
through the Mach number range by varying the
stagnation pressure,


NACA RM L50L07

AN EXPERIMENTAL STUDY OF MODERATE AND
HIGH SUBSONIC SPEEDS OF THE FLOW OVER AN
UNSWEPT WING IN CONJUNCTION WITH A FUSE-
LAGE. Richard T. Whitcomb. June 18, 1951. 35p.
diagrs., photos. (NACA RM L50L07) (Declassified
from Confidential, 8/31/54)

A relatively extensive study is presented of the pres-
sure distributions, wake measurements, and tuft pat-
terns obtained for an unswept, high-aspect-ratio,
tapered wing, in conjunction with a fuselage, at Mach
numbers up to 0.925.


NACA RM L50L12a

TABULATED PRESSURE COEFFICIENTS AND
AERODYNAMIC CHARACTERISTICS MEASURED
IN FLIGHT ON THE WING OF THE DOUGLAS
D-558-1 AIRPLANE THROUGHOUT THE NORMAL-
FORCE-COEFFICIENT RANGE AT MACH NUMBERS
OF 0.67, 0.74, 0.78, AND 0.82. Earl R. Keener,
James R. Peele and Julia B. Woodbridge.
January 29, 1951. 37p. diagrs., photos., 6 tabs.
(NACA RM L50L12a) (Declassified from Confidential,
8/31/54)

Tabulated pressure coefficients and aerodynamic
characteristics measured in flight on the right wing
of the Douglas D-558-I airplane throughout the
normal-force-coefficient range at Mach numbers of
0.67, 0.74, 0.78, and 0.82 are presented.


NACA RM L50L18

FLIGHT DETERMINATION OF THE DRAG AND
PRESSURE RECOVERY OF AN NACA 1-40-250
NOSE INLET AT MIACHf NUMBERS FROM 0.9 TO
1.8. R. I. Sears and C. F. Merlet. February 28,
1951. 32p. diagrs., photos., 2 tabs. (NACA
RM L50L18) (Declassified from Confidential,
8/3 1/54)

External-drag and pressure-recovery data are
presented for the NACA 1-40-250 nose inlet. The
tests were made using rocket-propelled models in
free flight at Mach numbers from 0.9 to 1.8. The
Reynolds number based on body diameter varied
from 4 to 10 x 106.


NACA
RESEARCH ABSTRACTS NO. 70

NACA RM L50L28

ERROR IN AIRSPEED MEASUREMENT DUE TO
STATIC-PRESSURE FIELD AHEAD OF THE WING
TIP OF A SWEPT-WING AIRPLANE MODEL AT
TRANSONIC SPEEDS. Edward C. B. Danforth and
Thomnas C. O'Bryan. March 1, 1951. 16p. diagrs.,
photo. (NACA RM L50L28) (Declassified from
Confidential, 8/31/54)

Contains measurements of static pressure taken at a
distance of 1 tip chord ahead of the wiing tip of a
model of a swept-wing fighter airplane for the con-
dition of near zero lift. The measurements made by
means of the NACA wing-flow method cover a range
of Mach numbers from 0.7 to 1.08. The influence of
the complete airplane configuration on the static
pressure is shown, together with the influences of
the wing and fuselage taken separately. Comparisons
with the linear theory are made where possible.

NACA RM L51A16

A COMPARISON OF TWO TECHNIQUES UTILIZING
ROCKET-PROPELLED VEHICLES FOR THE DE-
TERMINATION OF THE DAMPING-IN-ROLL
DERIVATIVE. David G. Stone and Carl A. SandahL.
May 3, 1951. 17lp. diagrs., photos. (NACA
RM L51A16) (D~eclassified from Confidential,
8/3 1/54)

Rocket-powered flight investigations have been con-
ducted for the purposes of comparing damping-in-
roll derivatives as obtained from the torque-nozzle
technique and the sting-mount technique. The
agreement in damping-in-roll values obtained by the
two techniques was good and showed the transonic
rolling behavior of wings to be more affected by
wing-dropping characteristics than by damping in
roll.


NACA RM L51A17

APPLICATION OF THEODORSEN'S PROPELLER
THEORY TO THE CALCULATION OF THE PER-
FORMANCE OF DUAL-ROTATING PROPELLERS.
Jean Gilman, Jr. March 15, 1951. 31p. diagrs.
(NACA RM L51A17) (Declaslsified from Confidential,
8/31/54)

Theodorsen's propeller theory is used to calculate
the performance of dual-rotating propellers having
nonideal load distributions. The ratio of the
spinner radius to the propeller radius is shown to
have a significant effect in evaluating the mass coef-
ficient used in making calculations for either per-
formance or design work. Sample performance
calculations are made and the results are compared
with experimental results for flight Mach nubes
varying from 0.53 to 0.00.


NACA RM L51A18

CORRELATION OF SUPERSONIC CONVECTIVE
HEAT-TRANSFER COEFFICIENTS FROM. MEAS-
UREMENTS OF THE SKIN TEMPERATURE OF A
PARABOLIC BODY OF REVOLUTION (NACA RM-10).
Leo T. Chauvin and Carlos A. deMoraes. March 7,
1951. 39p. diagrs., photo. 2 tabs. (N~ACA
RM L51A18) (Declassified from Confidential,
8/31i/54)








NACA
RESEARCH ABSTRACTS NO. 70

Free-flight tests at sutpersonic speeds have been
made to determine the local convective heat-transfer
coefficients, evaluated from measured skin tempera-
tures along the body of a rocket-propelled fin-
stabilized parabolic body of revolution. The Mach
number range covered was 1.02 to 2.48 and the
Reynolds number range was 3.18 x 106 to 1163.85 x 106
The experimental values are compared with the
results obtained from the V-2 research missile and
also writh several equations for heat transfer in a
turbulent boundary layer.



NACA RM L51A19

THE TORSIONAL DEFLECTIONS OF SEVERAL
PROPELLERS UNDER OPERATING CONDITIONS.
W. H.Gray and A. E.Allis. June 1951. 56p.
diagrs., photos., tab. (NACA RM L51A19) (Declas-
sified from Confidential, 8/31/54)

Presents measured and computed torsional deflec-
tions of the family of constant-chord solid aluminum-
alloy propeller blades used in a concurrent investi-
gation of blade-section pressure distributions. The
deflections were not negligible and varied with blade
design and operating condition. It was concluded
that blade torsional deflection may be computed with
sufficient accuracy and should be considered in the
design of thin propeller blades.


NACA RM L51A23

FLIGHT DETERMINATION OF THE EFFECTS OF
WING VORTEX GENERATORS ON THE AERO-
DYNAMIC CHARACTERISTICS OF THE DOUGLAS
D-558-I AIRPLANE. De E. Beeler, Donald R.
Bellman and John H. Griffith. August 14, 1951. 23p.
diags., photos., tab. (NACA RM L51A23) (Declas-
sified from Confidential, 8/31/54)

Tests were made to determine the effects of wing
vortex generators on the handling and buffe~ting char-
acteristics of the Douglas D-558-I airplane. M~eas-
urements of the chordwise pressure distribution over
one section of the wing, the total-head losses in a
portion of the wing wake, the total airplane drag, and
the buffeting and handling characteristics were made
with the basic configuration and with vortex genera-
tors of an arbitrary size, shape, and location
installed on the wing.



NACA RM L51B12

AVERAGE SKIN-FRICTION COEFFICIENTS FROM
BOUNDR-LAYER MEASUREMENTS IN FLIGHT
ON A PARABOLIC BODY OF REVOLUTION (NACA
RM-10) AT SUPERSONIC SPEEDS AND AT LARGE
REYNOLDS NUMB~ERS. Charles B. Rumsey and J.
Dan Loposer. March 7, 1951. 33p. diagrs., photo.
(NACA RM L51B12) (D~eclassified from Confidential,
8/31/54)

Boundary-layer measurements were made at two
stations on a rocket-powered, fin-stabilized, para-
bolic body of revolution of fineness ratio 12.2. Skin-
friction coefficients were determined over a Mach
numer range from 1.5 to 3.4 and over the Reynolds
number range, based on body length to the measure.
ment station, from 42 x 106 to 160 x 106. Theoret-


ical turbulent skin-friction coefficients for a flat
plate, with heat transfer, are also shown.



NACA RM L51Bl3

THE EFFECT OF END PLATES, END STRUTS,
AND DEPTH OF SUBMERGENCE ON THE CHAR-
ACTERISTICS OF A HYDROFOIL. Kennelth L.
Wadlin, Rudolph E. Fontana and Charles L. Shuford,
Jr. April 12, 1951. 84p. diagrs., photos. (NACA
RM L51B13) (Declassified from Confidential,
8/3 1/54)

An investigation was made in Langley tank No. 2, at
subcavitation speeds, of the effect of end plates and
end-mounted struts. Only small improvements
might be expected in the maxium lift-drag ratios
by the addition of end plates. End struts reduced
the lift-drag ratio attainable with a single strut. The
effect of end plates and end struts on effective aspect
ratio was in good agreement with theory. An ap-
proximate theoretical solution of the effect of depth
on the lift of a hydrofoil was developed.



NACA RM L51B28

NOTE ON FLUTTER OF A 600 DELTA WING
ENCOUNTERED AT LOW-SUPERSONIC SPEEDS
DURING THE FLIGHT OF A ROCKET-PROPELLED
MODEL. William T`. Lauten, Jr. and Grady L.
Mitcham. May 14, 1951. 21p. diagrs., photos.,
5 tabs. (NACA RM L51B28) (Declassified from
Confidential, 8/31/54)

Results of the flight test of a rocket-powered model
of a delta-wing (600 sweepback) body configuration
indicated a wing flutter at a Mach number of 1.11 and
a subsequent structural failure of a Mach number of
0.99. Sections of the flght time history and the
ground-test results of a duplicate wing for natural
frequencies of vibration, the structural influence
coefficients, and the mass, moment of inertia, and
center of gravity of streamwise strips of the wing
are presented.



NACA RM L51C07

EFFECT OF A DEFLECTABLE WIG-TIP CON-
TROL ON THE LOWY-8PEED LATERAL AND
LO)NGIUDINAL CHARACTERISTICS OF A LARGE-
SCALE. WING WITH THE: LEADING EDGE SWEPT
BACK 47.50. Roy H. Lange and Marvin P. Fink.
April 26, 1951. 41p. diagrs., photo., 2 tabs.
(NACA RM L51C07) (Declassified from Confidential,
8/31/54)

Results are presented of an investigation in the
Langley full-scale tunnel of the effect of a 20-
percent-semispan deflectable wing-tip control on the
low-speed lateral and longitudinal characteristics of
a wing with the leading edge swept back 47.50 and
circular-are-airfoil sections. The basic wing con-
figurations, the wing with drooped-nose flaps de-
flected 400, and the wing with drooped-nose and
semispan plain flaps deflected 400 were investigated
throughout. All the data are presented for a Reynolds
number of 4.3 x 106 and a Mach number of 0.07.








NACA
RESEARCH ABSTRACTS NO.70


NACA RM L51C26

AERODYNAMIC CHARACTERISTICS OF TAPERED
WINGS HAVING ASPECT RATIOS OF 4, 6, AND 8,
QUARTER-CHORD LINES SWEPT BACK 450, AND
NACA 631A012 AIRFOIL SECTIONS. TRANSONIC-
BUMP METHOD. Edward C. Polhamus and
Thomas J. King, Jr. June 13, 1951. 23p. diagrs.,
photos., 2 tabs. (NACA RM L51C26) (Declassified
from Confidential, 8/31/54)

This paper presents the results of an investigation
by the transonic-bump method of wings of aspect
ratios 8, 6, and 4 with the quarter-chord lines
swept back 450, and having NACA 631A012 airfoil
sections, parallel to the plane of symmetry. Lift,
drag, pitching-moment, and root bending-moment
characteristics are presented for the Mach number
range from 0.70 to 1.15.




NACA RM L51D18

FLIGHT INVESTIGATION OF THE LONGITUDINAL
STABILITY AND CONTROL CHARACTERISTICS OF
THE DOUGLAS D-558-I AIRPLANE (BUAERO
NO 37972) AT MACH NUMBERS UP TO 0.89. Melvin
Sadoff, William. S. Roden and John M. Eggleston.
June 1951. 26p. diagrs., photos., tab. (NACA
RM L51D18) (Declassified from Confidential,
8/3 1/54)

Results and analysis pertaining to the longitudinal
stability and control characteristics of the Douglas
D-558-I airplane (BuAero No. 37972) are presented.
The results indicated that large and rapid changes in
elevator deflection and force were required for bal-
ance above a Mach number of 0.84. Analysis indicat-
ed that a major part of these changes was due to a
loss in elevator effectiveness. A large increase in
the apparent stick-fixed stability parameter d8/dCN
was also noted due to a loss in elevator effectiveness
combined with an increase in airplane stability.




NACA RM L51D19

EFFECTS OF SPANWISE THICKNESS VARIATION
ON THE AERODYNAMIC CHARACTERISTICS OF
350 AND 450 SWEPTBACK WINGS OF ASPECT
RATIO 6. TRANSONIC-BUMP METHOD. William
D. Morrison, Jr. and Paul G. Fournier. July 1951.
38p. diagrs., photo. (NACA RM L51D19) (Declas-
sified from Confidential, 8/31/54)

The effects of taper-in-thickness on the aerodynamic
characteristics of wings having 350 and 450 of sweep-
back, aspect ratio 6, and taper ratio 0.6 have been
determined by the transonic-bump technique over a
Mach number range from 0.6 to 1.14. The results
of this investigation are compared with those obtain-
ed from wings of the same plan form but of various
constant section thickness ratios. Theoretical sub-
sonic and low supersonic calculations of lift-curve
slope, aerodynamic center, and lateral center-of-
pressure location were determined.


NACA RM L51D23

THEORETICAL INVESTIGATION OF AN AUTO-
MATIC CONTROL SYSTEM WITH PRIARY SENSI-
TIVITY TO NORMAL ACCELERATIONS AS USED
TO CONTROL A SUPERSONIC CANARD MISSILE
CONFIGURATION. Ernest C. Seaberg and Earl F.
Smith. July 1951. 48p. diagrs., photo., 3 tabs.
(NACA RM LS1D23) (Declassified from Confidential,
8/3 1/54)

An analysis has been made to determine the possibil-
ities of using an autopilot primarily sensitive to lin-
ear accelerations for longitudinal control of a spe-
cific supersonic canard airframe. Essentially, the
control system combines the use of a linear accel-
erometer and an integrating servomotor to obtain
desired normal accelerations of the missile. The
analysis is based largely on comparisons of normal
acceleration transient responses obtained for var-
ious conditions of Mach number, altitude, static
margin, and rate-of-pitch feedback. Because the
acceleration control system has no directional space
reference of its own, it is believed that its primary
usefulness is in conjunction with a homing seeker or
with a guidance system which will provide a direc-
tional space reference.


NACA RM L51D24

PRESSURE-DISTRIBUT~ION MEASUREMENTS OVER
A 450 SWEPTBACK WING AT TRANSONIC SPEEDS
BY THE NACA WING-FLOW METHOD. Edward
C. B. Danforth and Thomas C. O'Bryan. Jun 1951.
42p. diagrs., photos. (NACA RM L51D24) (Declas-
sified from Confidential, 8/31/54)

Pressure distributions have been obtained over the
chord of an untapered, 450 sweptback wing model of
3.5 aspect ratio with 2-inch-chord NACA 65-210 air-
foil sections normal, to the leading edge at four sta-
tions along the span. The measurements, made by
the NACA wing-flow method, covered Mach numbers
from 0.7 to 1.1 and angles of attack from -lo to 40 at
a nominal Reynolds number of 0.6 x 106 based on the
chord. Comparisons are made with force tests and
theory where possible.


NACA RM L51D24a

WING-FLOW INVESTIGATION OF THE CHARAC-
TERISTICS OF SEVEN UNSWEPT, UNTAPERED
AIRFOILS3 OF ASPECT RATIO 8.0. Harold L. Crane
and James J. Adams. June 13, 1951. 54p. diagrs.,
photo. (NACA RM L51D24a) (Declassified from
Confidential, 8/31/54)

Measurements were made of normal force, chord
force, and pitching moment at angles of attack from
-6o to 140 and Mach numbers of 0.65 to 1.08, for
seven airfoil sections. The sections included were
NACA 65-010, 65-210, 836D110, 847B110, a 10-
percent-thick airfoil with 3-to-1 elliptical nose
faired into a straight-sided afterbody, an airfoil
having the same thickness distribution and a reflex-
camber line obtained by subtracting an NACA 240
from, an NACA 420 mean line, and an airfoil hain
the thickness distribution of the NACA 65-010 and
the reflex-camber line.


NACA-I.angley 9-21-54 4M


UN~~~lIVERSIYO LRD




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