| 1. |
The temperature of unheated
bodies in a high-speed gas stream (1941) by Eckert, E.,
Weise, W. [21 pages; 1.1 MB] |
|
Abstract: The present report deals with
temperature measurements on cylinders of 0.2 to 3 millimeters diameter
in longitudinal and transverse air flow at speeds of 100 to 300 meters
per second. Within the explored test range, that is, the probable
laminar boundary layer region, the temperature of the cylinders in
axial flow is practically independent of the speed and in good
agreement with Pohlhausen's theoretical values; Whereas, in transverse
flow, cylinders of certain diameter manifest a close relationship with
speed, the ratio of the temperature above the air of the body to the
adiabatic stagnation temperature decreases with rising speed and then
rises again from a Mach number of 0.6. The importance of this "specific
temperature" of the body for heat-transfer studies at high speed is
discussed. |
| 2. |
An investigation of the drag of
windshields in the 8-foot high-speed wind tunnel (1942)
by Robinson, Russell G Delano, James B [13 pages; 0.7 MB] |
|
Abstract: Report presents the results of tests
made to determine the drag of closed-cockpit and transport-type
windshields. The tests were made at speeds corresponding to a Mach
number range of approximately 0.25 to 0.58 in the NACA 8-foot
high-speed wind tunnel. This speed range corresponds to a test Reynolds
number range of 2,510,000 to 4,830,000 based on the mean aerodynamic
chord of the full-span model (17.29 in.). The shapes of the windshield
proper, the hood, and the tail fairing were systematically varied to
include common types and refined design. |
| 3. |
Tests of airfoils designed to
delay the compressibility burble (1943) by Stack, John
[14 pages; 0.9 MB] |
|
Abstract: Fundamental investigations of
compressibility phenomena for airfoils have shown that serious adverse
changes of aerodynamic characteristics occur as the local speed over
the surface exceeds the local speed of sound. These adverse changes
have been delayed to higher free-stream speeds by development of
suitable airfoil shapes. The method of deriving such airfoil shapes is
described, and aerodynamic data for a wide range of Mach numbers
obtained from tests of these airfoils in the Langley 24-inch high-speed
tunnel are presented. These airfoils, designated the NACA 16-series,
have increased critical Mach number. The same methods by which these
airfoils have been developed are applicable to other airplane
components. |
| 4. |
The flow of a compressible
fluid past a curved surface (1943) by Kaplan, Carl [23
pages; 1.2 MB] |
|
Abstract: An iteration method is employed to
obtain the flow of a compressible fluid past a curved surface. The
first approximation which leads to the Prandtl-Glauert rule, is based
on the assumption that the flow differs but little from a pure
translation. The iteration process then consists in improving this
first approximation in order that it will apply to a flow differing
from pure translatory motion to a greater degree. The method fails when
the Mach number of the undisturbed stream reaches unity but permits a
transition from subsonic to supersonic conditions without the
appearance of a compression shock. The limiting value at which
potential flow no longer exits is indicated by the apparent divergence
of the power series representing the velocity of the fluid at the
surface of the solid boundary. |
| 5. |
A simplified chart for
determining Mach number and true airspeed from airspeed-indicator
readings (March 1943) by Ritchie, Virgil S [9 pages; 0.4
MB] |
|
Abstract: No Abstract Available |
| 6. |
The relation between spanwise
variations in the critical Mach number and spanwise load distributions
(December 1944) by Richard T. Whitcomb [9 pages; 0.7 MB] |
|
Abstract: Data are presented to show the changes
that occur in the spanwise load distributions on wings when the
critical Mach number is exceeded. These data indicate that the
magnitude of the spanwise variation in the critical Mach numbers of the
sections. Means of reducing the magnitudes of such changes are
considered. |
| 7. |
On the flow of a compressible
fluid by the hodograph method I : unification and extension of
present-day results (1944) by Garrick, I E Kaplan, Carl
[24 pages; 1.3 MB] |
|
Abstract: Elementary basic solutions of the
equations of motion of a compressible fluid in the hodograph variables
are developed and used to provide a basis for comparison, in the form
of velocity correction formulas, of corresponding compressible and
incompressible flows. The known approximate results of Chaplygin, Von
Karman and Tsien, Temple and Yarwood, and Prandtl and Glauert are
unified by means of the analysis of the present paper. Two new types of
approximations, obtained from the basic solutions, are introduced; they
possess certain desirable features of the other approximations and
appear preferable as a basis for extrapolation into the range of high
stream Mach numbers and large disturbances to the main stream. Tables
and figures giving velocity and pressure-coefficient correction factors
are included in order to facilitate the practical application of the
results. |
| 8. |
Compressible potential flow
with circulation about a circular cylinder (1944) by
Heaslet, Max A [9 pages; 0.5 MB] |
|
Abstract: The potential function for flow, with
circulation, of a compressible fluid about a circular cylinder is
obtained in series form including terms of the orders of m(4) where m
is the Mach number of the free stream. The resulting equations are used
to obtain pressure coefficients as a function of Mach number at a point
on the surface of the cylinder for different values of circulation. The
coefficients derived are compared with the Glauert-Prandtl and
Karman-Tsien approximations which are functions of the pressure
coefficients of an incompressible fluid. For the cases considered, the
values of the pressure coefficients computed from the theory were found
to be somewhere between the two approximations, the first
underestimating and the second overestimating it. |
| 9. |
Experiments on drag of
revolving disks, cylinders, and streamline rods at high speeds
(1944) by Theodorsen, Theodore Regier, Arthur [18 pages; 1 MB] |
|
Abstract: An experimental investigation concerned
primarily with the extension of test data on the drag of revolving
disks, cylinders, and streamline rods to high Mach numbers and Reynolds
numbers is presented. |
| 10. |
A method for the calculation of
external lift, moment, and pressure drag of slender open-nose bodies of
revolution at supersonic speeds (1945) by Brown, Clinton
E Parker, Hermon M [8 pages; 0.5 MB] |
|
Abstract: An approximate method is presented for
the calculation of the external lift, moment, and pressure drag of
slender open-nose bodies of revolution of supersonic speeds. The lift,
moment, and pressure drag of a typical ram-jet body shape are
calculated at Mach numbers 1.45, 1.60, 1.75, and 3.00; and the lift and
moment results are compared with available experimental data. The
agreement of the calculated lift and moment data with the experimental
data is excellent. The pressure-drag comparison was not presented
because of the uncertainty of the amount of skin-friction drag present
in the experimental results. |
| 11. |
A systematic investigation of
pressure distributions at high speeds over five representative NACA
low-drag and conventional airfoil sections (1945) by
Graham, Donald J Nitzberg, Gerald E Olson, Robert N [68 pages; 4.6 MB] |
|
Abstract: Pressure distributions determined from
high-speed wind-tunnel tests are presented for five NACA airfoil
sections representative of both low-drag and conventional types.
Section characteristics of lift, drag, and quarter-chord pitching
moment are presented along with the measured pressure distributions for
the NACA 65sub2-215 (a=0.5), 66sub2-215 (a=0.6), 0015, 23015, and 4415
airfoils for Mach numbers up to approximately 0.85. A critical study is
made of the airfoil pressure distributions in an attempt to formulate a
set of general criteria for defining the character of high speed flows
over typical airfoil shapes. Comparisons are made of the relative
characteristics of the low-drag and conventional airfoils investigated
insofar as they would influence the high-speed performance and the
high-speed stability and control characteristics of airplanes employing
these wing sections. |
| 12. |
Graphical and analytical
methods for the determination of a flow of a compressible fluid around
an obstacle (1945) by Bergman, Stefan (Brown University,
Providence, R.I) [38 pages; 1 MB] |
|
Abstract: Chaplygin introduced the hodograph
method in the theory of compressible fluid flows and developed a method
for constructing stream functions of such flows. This method, which has
been extensively used in investigation of compressible fluid flows, is
limited in certain respects. The expression for the stream function
obtained in this manner can represent only certain types of flow
patterns. In general, flow patterns obtained in this way cannot
represent the whole flow around an obstacle, but only a part of such a
flow, and therefore several expressions are needed in order to obtain
the whole flow. On the other hand, in many instances it is important to
have a single expression representing the whole flow. Recently Von
Karman and Tsien constructed more general types of stream functions,
but only by replacing the true pressure density relation by the linear
pressure-specific volume relation so that their method is essential
limited to flows the maximum Mach number of which is not too large. In
a companion report the author derived a new formula for stream
functions based on the true pressure density relation. It is not
subject to the limitations of the Chaplygin method. In the present
report this formula is employed to construct two-dimensional subsonic
compressible fluid flows around a body similar in shape to a given
symmetric obstacle. The methods described in the report are illustrated
by numerical examples. |
| 13. |
On the circulatory subsonic
flow of a compressible fluid past a circular cylinder (1945)
by Bers, Lipman (Brown University, Providence, R.I) [43 pages; 1 MB] |
|
Abstract: The circulatory subsonic flow around an
infinite circular cylinder is computed using the linearized
pressure-volume relation, by a method developed in a previous report.
Formulas and graphs are given for the velocity and pressure
distributions, the circulation, the lift, and the dependence of the
critical Mach number upon the position of the stagnation point. |
| 14. |
Effect of propeller-axis angle
of attack on thrust distribution over the propeller disk in relation to
wake-survey measurement of thrust (December 1945) by
Pendley, Robert E [34 pages; 1.1 MB] |
|
Abstract: Tests were made to investigate the
variation of thrust distribution over the propeller disk with angle of
pitch of the propeller thrust axis and to determine the disposition and
the minimum number of rakes necessary to measure the propeller thrust.
The tests were made at a low Mach number for a low and a high blade
angle with the propeller operating at three small angles of pitch, and
some of the tests were repeated at a higher Mach number. The data
obtained show that, for small angles of pitch, large changes occur in
the energy distribution in the wake which prohibit the use of a single
survey rake for thrust measurement in flight tests and limit the use of
a single rake in wind-tunnel tests. Under certain conditions, the
energy distribution in the wake took on a symmetrical form and two
diametrically opposed survey rakes were shown to be satisfactory for
obtaining propeller thrust. (author) |
| 15. |
Properties of low-aspect-ratio
pointed wings at speeds below and above the speed of sound
(1946) by Jones, Robert T [18 pages; 0.6 MB] |
|
Abstract: Low-aspect-ratio wings having pointed
plan forms are treated on the assumption that the flow potentials in
planes at right angles to the long axis of the airfoils are similar to
the corresponding two-dimensional potentials. For the limiting case of
small angles of attack and low aspect ratios the theory brings out the
following significant properties: (1) The lift of a slender, pointed
airfoil moving in the direction of its long axis depends on the
increase in width of the sections in a downstream direction. Sections
behind the section of maximum width develop no lift. (2) The spanwise
loading of such an airfoil is independent of the plan form and
approaches the distribution giving a minimum induced drag. (3) The lift
distribution of a pointed airfoil travelling point-foremost is
relatively unaffected by the compressibility of the air below or above
the speed of sound. A best of a triangular airfoil at a Mach number of
1.75 verified the theoretical values of lift and center of pressure. |
| 16. |
Two-dimensional irrotational
mixed subsonic and supersonic flow of a compressible fluid and the
upper critical Mach number (May 1946) by Tsien,
Hsue-Shen (California Institute of Technology) Kuo, Yung-Huai
(California Institute of Technology) [138 pages; 4 MB] |
|
Abstract: No Abstract Available |
| 17. |
Investigation of the
characteristics of a high-aspect-ratio wing in the Langley 8-foot
high-speed tunnel (Aug 1946) by Richard T. Whitcomb [78
pages; 2.6 MB] |
|
Abstract: An investigation of the characteristics
of a wing with an aspect ratio of 9.0 and and NACA 65-210 airfoil
section has been made at Mach numbers up to 0.925. The wing tested has
a taper ratio of 2.5:1.0, no twist, dihedral, or sweepback, and a 20
percent chord 37.5 percent semispan plain ailerons. |
| 18. |
Aerodynamic characteristics
including scale effect of several wings and bodies alone and in
combination at a Mach number of 1.53 (December 20, 1946)
by Van Dyke, Milton D [85 pages; 2.6 MB] |
|
Abstract: No Abstract Available |
| 19. |
Effect of Mach and Reynolds
numbers on maximum lift coefficient (March 28, 1946) by
Spreiter, John R Steffen, Paul J [40 pages; 1.5 MB] |
|
Abstract: No Abstract Available |
| 20. |
Note on the theorems of
Bjerknes and Crocco (May 1946) by Theodorsen, Theodore
[5 pages; 0 MB] |
|
Abstract: The theorems of Bjerknes and Crocco are
of great interest in the theory of flow around airfoils at Mach numbers
near and above unity. A brief note shows how both theorems are
developed by short vector transformations. |
| 706. |
An experimental study of the
lift and pressure distribution on a double-wedge profile at Mach
numbers near shock attachment (Jul 1954) by Walter G.
Vincenti, Duane W. Dugan, E. Ray Phelps [44 pages; 1.6 MB] |
|
Abstract: An account is given of wind-tunnel
measurements at low supersonic speeds of the pressure distribution on a
doubly symmetrical double-wedge profile of approximately 8-percent
thickness. The results cover the Mach number range form 1.166 to 1.377,
which brackets the value (1.221) given by exact inviscid theory for
attachment of the shock wave to the leading edge at zero angle of
attack. |
| 707. |
Investigation of distributed
surface roughness on a body of revolution at a Mach number of 1.61
(Jun 1954) by K. R. Czarnecki, Ross B. Robinson, John H. Hilton,
Jr. [36 pages; 1 MB] |
|
Abstract: An investigation has been made of the
effects of distributed surface roughness, consisting of lathe-tool
marks, on the skin-friction drag of a body of revolution at a Mach
number of 1.61. The tests were made on ogive-cylinders at zero angle of
attack over a roughness range from 23 to 480 microinches root mean
square over Reynolds number range from 2.5 X 10(exp 6) to 37 X 10(exp
6). |
| 708. |
Experimental investigation of
temperature recovery factors on a 10 degree cone at angle of attack at
a Mach number of 3.12 (Jul 1954) by John R. Jack, Barry
Moskowitz [16 pages; 0.4 MB] |
|
Abstract: Temperature recovery factors on a
thin-walled, metal, 10 degree included angle cone were obtained at a
Mach number of 3.12 over a range of angles of attack from 0 to 10
degrees and for Reynolds numbers per foot from 1.5 X 10 (exp 6) to 8 X
10(exp 6). |
| 709. |
Investigation of Mach number
changes obtained by discharging high-pressure pulse through wind tunnel
operating supersonically (Aug 1954) by Rudolph C.
Haefeli, Harry Bernstein [15 pages; 0.4 MB] |
|
Abstract: A series of tests was performed to
obtain an indication of the transient-flow phenomena caused by
discharging a chamber of high-pressure gas into a wind tunnel operating
supersonically. For the configurations tested, two types of gust were
obtained. One had a maximum Mach number with a practically zero time
duration. The other had a maximum Mach number with a finite time
duration depending on the specific geometry. Such configurations are
applicable as supersonic longitudinal-gust tunnels. |
| 748. |
A summary of information on
support interference at transonic and supersonic speeds
(January 12, 1954) by Love, Eugene S [28 pages; 0.8 MB] |
|
Abstract: An experimental investigation was
performed to determine the effect on base and forebody pressures of
using a sting modified with varying length splitter plates and fins
instead of a conventional sting to support a cone-cylinder body of
revolution. The investigation was conducted at a Mach number of 3.12
for a Reynolds number range of 2 x 10 to the 6th power to 14 x 10 to
the 6th power and for an angle of attack range of 0 degrees to 9
degrees. For Reynolds numbers of 8 x 10 to the 6th power and 14 x 10 to
the 6th power there was a negligible effect of the splitter plate
modification on the base pressure, and at Reynolds number of 2 x 10 to
the 6th power there was a small effect. Positioning the leading edge of
the splitter plate at or ahead of the base made no appreciable change
in the influence of the modifications on base pressure at a Reynolds
number of 14 x 10 to the 6th power. With the fin-type modification
there was a small increase in base pressure. |
| 749. |
An air-flow-direction pickup
suitable for telemetering use on pilotless Aircraft (March 10,
1954) by Ikard, Wallace L [27 pages; 1 MB] |
|
Abstract: A free-swiveling vane-type pickup for
measuring air flow direction in both the angle-of-attack and
angle-of-sideslip directions is described. The device, which is
intended to telemeter flow direction from pilotless aircraft, has
variable-inductance outputs suitable for use in the 100 to 200 kcps
subcarrier frequency range of the NACA FM-AM telemetering system.
Preliminary test results indicate that it can also be adapted for use
with the audio subcarrier frequencies of the Research and Development
Board standard FM-FM telemetering system. Test results are presented
which indicate that the pickup is aerodynamically stable and has an
accuracy, obtained from a bench calibration, of better than 0.3 degrees
under conditions including acceleration up to 20g in any direction,
Mach numbers from 0.5 to 2.8, and dynamic pressures up to at least 65
psi. Equations and curves which can be used to obtain flow direction at
the center of gravity of a maneuvering model are presented. |
| 750. |
Flight investigation of the
rolling effectiveness of fingered semaphore spoilers on a tapered 45
sweptback wing between Mach numbers 0.6 and 1.3 (January 14,
1954) by Church, James D [29 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 751. |
Experimental convective heat
transfer to a 4-inch and 6-inch hemisphere at Mach numbers from 1.62 to
3.04 (February 03, 1954) by Chauvin, Leo T Maloney,
Joseph P [20 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 752. |
Free-flight measurements of the
rolling effectiveness and drag of trailing-edge spoilers on a tapered
sweptback wing at Mach numbers 0.6 and 1.4 (February 18, 1954)
by Schult, Eugene, D Fields, E M [15 pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 753. |
Aerodynamic characteristics of
a full-span trailing-edge control on a 60 degree delta wing with and
without a spoiler at Mach number 1.61 (March 10, 1954)
by Lord, Douglas R Czarnecki, K R [51 pages; 1.3 MB] |
|
Abstract: No Abstract Available |
| 754. |
Effects of sweep and thickness
on the static longitudinal aerodynamic characteristics of a series of
thin, low-aspect-ratio, highly tapered wings at transonic speeds :
transonic-bump method (April 08, 1954) by Fournier, Paul
G Few, Albert G , Jr [108 pages; 3.2 MB] |
|
Abstract: An investigation by the transonic-bump
technique of the static longitudinal aerodynamic characteristics of a
series of thin, low-aspect-ratio, highly tapered wings has been made in
the Langley high-speed 7- by 10-foot tunnel. The Mach number range
extended from about 0.60 to 1.18, with corresponding Reynolds numbers
ranging from about 0.75 x 10(6) to 0.95 x 10(6). The angle of attack
range was from -10 degrees to approximately 32 degrees.The effects on
drag and lift-drag ratio of a variation in sweep angle from -14.03
degrees to 45 degrees with respect to the quarter-chord line for wings
of 3-percent-chord thickness was found to be small in comparison to the
effects of a variation in thickness from 2 percent chord to 4.5 percent
chord for wings with 14.03 degree sweepback. For the range of variables
considered, variations in plan form were considerably more important
with regard to longitudinal stability characteristics than the
variations in thickness. For the series of basic wings having an aspect
ratio of 4, the most hearly linear pitching-moment characteristics were
obtained with 26.57 degree of sweepback of the quarter-chord line.
However, for the modified series of wings (obtained by clipping the
tips of the original wings parallel to the plane of symmetry to give an
aspect ratio of 3 and a taper ratio of 0.143), the most nearly linear
pitching-moment characteristics were obtained with 36.87 degrees of
sweepback. By decreasing the thickness-to-chord ratios from 0.03 to
0.02, a large increase in lift-curve slope was obtained for both the
basic and modified wings. All of the wings of both series had fairly
large inward shifts of the lateral center-of-pressure location
(indicative of tip stalling) with increasing lift coefficient, except
those wings having minimum sweepback angles. |
| 755. |
Rocket-powered model
investigation of lift, drag, and stability of a body-tail configuration
at Mach numbers from 0.8 to 2.3 and angles of attack between plus or
minus 6.5 degrees (April 15, 1954) by Gillespie, Warren,
Jr Dietz, Albert E [43 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 756. |
Measurements and predictions of
flow conditions on a two-dimensional base separating a Mach number 3.36
jet and a Mach number 1.55 outer stream (May 07, 1954)
by Coletti, Donald E [58 pages; 1.9 MB] |
|
Abstract: No Abstract Available |
| 757. |
Rocket-powered-model
investigation of the hinge-moment and normal-force characteristics of a
half-diamond tip control on a 60 degree sweptback diamond wind between
Mach numbers of 0.5 and 1.3 (April 26, 1954) by Church,
James D [32 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 758. |
Experimental effects of
propulsive jets and afterbody configurations on the zero-lift drag of
bodies of revolution at a Mach number of 1.59 (April 22, 1954)
by De Moraes, Carlos A Nowitzky, Albin M [34 pages; 1 MB] |
|
Abstract: The present investigation was made at a
free-stream Mach number of 1.59 to compare the afterbody drags to a
series of conical boattailed models at zero angle of attack. Afterbody
drags were obtained for both the power-off and the power-on conditions.
Power-on drags were obtained as a function of afterbody fineness ratio,
jet pressure ratio and divergence, and jet Mach number. |
| 759. |
A wind-tunnel investigation at
high subsonic speeds of the lateral control characteristics of various
plain spoiler configurations on a 3-percent-thick 60 degree delta wing
(May 26, 1954) by Wiley, Harleth G [47 pages; 1.9 MB] |
|
Abstract: Results are presented of wind-tunnel
investigations at Mach numbers of 0.60 to 0.94 and angles of attack of
-2 degrees to about 24 degrees to determine the lateral control
characteristics of spoilers with various wing chord-wise and spanwise
locations and spoiler spans and deflections on thin 60 degree delta
wing of NACA 65a003 airfoil section parallel to free stream. |
| 760. |
Drag and heat transfer on a
parabolic body of revolution (NACA RM-10) in free flight to Mach number
2 with both constant and varying Reynolds number and heating effects on
turbulent skin friction (June 17, 1954) by Maloney,
Joseph P [36 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 761. |
A preliminary investigation of
the pressure recovery of several two-dimensional supersonic inlets at
Mach number of 2.01 (June 23, 1954) by Comenzo, Raymond
J [30 pages; 0.7 MB] |
|
Abstract: No Abstract Available |
| 762. |
Aerodynamic characteristics of
several flap-type trailing-edge controls on a trapezoidal wing at Mach
numbers of 1.61 and 2.01 (June 14, 1954) by Lord,
Douglas R Czarnecki, K R [69 pages; 1.7 MB] |
|
Abstract: No Abstract Available |
| 763. |
Flight investigation of an
aileron and a spoiler on a wing of the X-3 airplane plan form at Mach
numbers from 0.5 to 1.6 (June 18, 1954) by English,
Roland D [17 pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 764. |
Flight investigation to
determine lift and drag characteristics of a canard ram-jet missile
configuration in the Mach number range of 0.8 to 2.0 (June 17,
1954) by Gammal, Abraham A Kennedy, Thomas L [21 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 765. |
Low-amplitude damping-in-pitch
characteristics of tailless delta-wing-body combinations at Mach
numbers from 0.80 to 1.35 as obtained with rocket-powered models
(June 24, 1954) by D'Aiutolo, Charles T [35 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 766. |
Normal force, center of
pressure, and zero lift drag of several ballistic-type missiles at Mach
numbers of 4.05 (July 06, 1954) by Ulmann, Edward F
Dunning, Robert W [30 pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 767. |
An investigation of the effects
of jet exhaust and Reynolds number upon the flow over the vertical
stabilizer and rudder of the Douglas D-558-II research airplane at Mach
numbers of 1.62, 1.93, and 2.41 (June 17, 1954) by
Grigsby, Carl E [40 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 768. |
Effect of wing flexibility on
the damping roll of a notched delta-wing body combination between Mach
numbers 0.6 and approximately 2.2 as determined with rocket-propelled
models (June 18, 1954) by Bland, William M , Jr [21
pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 769. |
An experimental investigation
of two-dimensional, supersonic cascade-type inlets at a Mach number of
3.11 (August 25, 1954) by Offenhartz, Edward [30 pages;
1 MB] |
|
Abstract: No Abstract Available |
| 770. |
Investigation of a canard
missile configuration (NACA RM-4) in the Langley 9-inch supersonic
tunnel at Mach numbers of 1.62 and 1.93 (June 24, 1954)
by Grigsby, Carl E [25 pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 771. |
Aerodynamic characteristics of
several tip controls on a 60 degree wing at a Mach number of 1.61
(August 05, 1954) by Lord, Douglas R Czarnecki, K R [45 pages;
0.9 MB] |
|
Abstract: No Abstract Available |
| 772. |
Wind-tunnel investigation at a
Mach number of 2.01 of the aerodynamic characteristics in combined
pitch and sideslip of some canard-type missiles having cruciform wings
and canard surfaces with 70 degree delta plan forms (August
23, 1954) by Spearman, M Leroy CORNELIUS DRIVER [122 pages; 2.8
MB] |
|
Abstract: No Abstract Available |
| 773. |
A method for increasing the
effectiveness of stabilizing surfaces at high supersonic Mach numbers
(August 03, 1954) by Mclellan, Charles H [15 pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 774. |
Investigation of the effect of
balancing tabs on the hinge-moment characteristics of a trailing-edge
flap-type control on a trapezoidal wing at a Mach number of 1.61
(August 05, 1954) by Driver, Cornelius Lord, Douglas R [24 pages;
0.6 MB] |
|
Abstract: No Abstract Available |
| 775. |
Aerodynamic characteristics at
Mach number of 2.01 of two cruciform missile configurations having 70
degree delta wings with length-diameter ratios of 14.8 and 17.7 with
several canard controls (August 30, 1954) by Spearman, M
Leroy Robinson, Ross B [33 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 776. |
Flight determination of the
drag of conical-shock nose inlets with various cowling shapes and axial
positions of the center body at Mach numbers from 0.8 to 2.0
(September 10, 1954) by Merlet, Charles, F Putland, Leonard W [42
pages; 1 MB] |
|
Abstract: No Abstract Available |
| 777. |
Effect of yaw and angle of
attack pressure recovery and mass-flow characteristics of a rectangular
supersonic scoop inlet at a Mach number of 2.71 (September 10,
1954) by Comenzo, Raymond J Mackley, Ernest A [21 pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 778. |
Drag measurements on a
1/6-scale, finless, sting-mounted NACA RM-10 missile in flight at Mach
numbers from 1.1 to 4.04 showing some Reynolds number and heating
effects (October 27, 1954) by Piland, Robert O [22
pages; 1 MB] |
|
Abstract: No Abstract Available |
| 779. |
The effect of a change in
airfoil section on the hinge-moment characteristics of a half-delta tip
control with a 60 degree sweep angle at a Mach number of 6.9
(October 15, 1954) by Fetterman, David E Ridyard, Herbert W [16
pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 780. |
Free-flight measurements of the
rolling effectiveness and operating characteristics of a
bellows-actuated split-flap aileron on a 60 degree delta wing at Mach
numbers between 0.8 and 1.8 (October 18, 1954) by
Schult, Eugene D [35 pages; 1.3 MB] |
|
Abstract: No Abstract Available |
| 781. |
An investigation of a
supersonic aircraft configuration having a tapered wing with
circular-arc sections and 40 degree sweepback : aerodynamic
characteristics of the configuration equipped with a canard control
surface at a Mach number of 1.89 (October 18, 1954) by
Spearman, M Leroy Plazzo, Edward B [24 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 782. |
An investigation of the
characteristics of the NACA RM-10 (with and without fins) in the
Langley 11-inch hypersonic tunnel at a Mach number of 6.9
(November 26, 1954) by Macauley, William D Feller, William V [41
pages; 2 MB] |
|
Abstract: No Abstract Available |
| 783. |
Low-amplitude damping-in pitch
characteristics of four tailless swept wing-body combinations at Mach
numbers from 0.85 to 1.30 as obtained with rocket-powered models
(November 24, 1954) by D'Aiutolo, Charles T [35 pages; 1.3 MB] |
|
Abstract: No Abstract Available |
| 784. |
Investigation at supersonic
speeds of the effect of jet Mach number and divergence angle of the
nozzle upon the pressure of the base annulus of a body of revolution
(December 17, 1954) by Bromm, August F O'Donnell, Robert M [25
pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 785. |
Experimental investigation of
effects of primary jet flow and secondary flow through a zero-length
ejector on base and boattail pressures of a body of revolution at
free-stream Mach numbers of 1.62, 1.93, and 2.41 (December 06,
1954) by O'Donnell, Robert M Mcdearmon, Russell W [42 pages; 1.3
MB] |
|
Abstract: An investigation was made at
free-stream Mach numbers of 1.62, 1.93, and 2.41 to determine the
effects of a primary jet and secondary air flow on the base pressure
and pressures acting over the boattailsurface of a body of revolution
for two secondary discharge areas. The Mach numbers of the primary
nozzles were 1 and 3.23 with the secondary mass flow being varied from
0 to 10 percent of the primary mass flow. The ratio of jet stagnation
temperature to tunnel stagnation temperature was about 0.96. The
Reynolds number range of the investigation was from 2.1 x 10(6) to 2.9
x 10(6)based on body length. All testing was conducted with a turbulent
boundary layer along the model. This report presents results obtained
with zero-length ejector and covers jet static-pressure ratios from the
jet-off condition to a maximum of about 128 for the sonic nozzle and to
a maximum of about 9 for the supersonic nozzle. |
| 786. |
Zero-lift drag of several
conical and blunt nose shapes obtained in free flight at Mach numbers
of 0.7 to 1.3 (March 23, 1954) by Piland, Robert O
Putland, Leonard W [16 pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 787. |
Effects of two spinner shapes
on the pressure recovery in an NACA 1-series D-type cowl behind a
three-blade propeller at Mach numbers up to 0.80 (March 19,
1954) by Reynolds, Robert M Molk, Ashley J [36 pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 788. |
An experimental investigation
of the flutter of several wings of varying aspect ratio, density, and
thickness ratio at Mach numbers from 0.60 to 1.10 (April 07,
1954) by Herrera, Raymond Barnes, Robert H [41 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 789. |
The effect of lip shape on a
nose-inlet installation at Mach numbers from 0 to 1.5 and a method for
optimizing engine-inlet combinations (May 07, 1954) by
Mossman, Emmet A Anderson, Warren E [50 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 790. |
Investigation of the normal
force accompanying thrust-axis inclination of the NACA
1.167-(0)(03)-058 and the NACA 1.167-(0)(05)-058 three-blade propellers
at forward Mach numbers to 0.90 (June 23, 1954) by
Demele, Fred A Otey, William R [33 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 791. |
A comparison of the
longitudinal aerodynamic characteristics at mach numbers up to 0.94 of
swept back wings having NACA 4-digit or NACA 64A thickness
distributions (August 23, 1954) by Sutton, Fred B
Dickson, Jerald K [69 pages; 2 MB] |
|
Abstract: No Abstract Available |
| 792. |
An experimental investigation
of reduction in transonic drag rise at zero lift by the addition of
volume to the fuselage of a wing-body-tail configuration and a
comparison with theory (August 18, 1954) by Holdaway,
George H [37 pages; 0.9 MB] |
|
Abstract: An experimental investigation was made
by the free-fall recoverable-model technique to assess at zero lift the
possibilities of reducing the drag-rise coefficients of a
wing-body-cruciform-tail combination by adding volume to the fuselage.
The basic features of the test model were an unswept aspect-ratio-3.1
thin wing, a fineness-ratio-12.4 fuselage, and four 45 degrees
sweptback tail surfaces. The tests covered a Mach number range of 0.84
to 1.15 with Reynolds numbers of 6.000.000 to 14,000,000, based on the
wing mean aerodynamic chord. Considerable reduction in drag-rise
coefficient was effected for several different modifications by the
addition of properly distributed volume to the fuselage. In one
instance, a reduction in drag coefficient was obtained by adding a
volume which was almost four times the exposed wing volume. The
computation method presented in NACA RM A53H17 generally predicted the
supersonic drag-rise coefficients for each modification within 20
percent of the experimental values. As in the above-mentioned report,
the predictions at a Mach number of one were not accurate. The changes
in drag-rise coefficients resulting from the modifications were
generally predicted with better accuracy than the values of drag-rise
coefficients. |
| 793. |
Investigation of the NACA
4-(5)(05)-037 six- and eight-blade, dual-rotation propellers at
positive and negative thrust at Mach numbers up to 0.90, including some
aerodynamic characteristics of the NACA 4-(5)(05)-041 two- and
four-blade, single-rotat (October 08, 1954) by Reynolds,
Robert M Walker, John H [150 pages; 10.9 MB] |
|
Abstract: No Abstract Available |
| 794. |
Effect of rotation of a NACA
1-series E-type cowling on the internal flow and force characteristics
of the cowling at Mach numbers up to 0.84 and at an angle of attack of
0 degrees (October 27, 1954) by Sammonds, Robert I
Reynolds, Robert M [56 pages; 1.6 MB] |
|
Abstract: No Abstract Available |
| 795. |
Longitudinal aerodynamic
characteristics to large angles of attack of a cruciform missile
configuration at a Mach number of 2 (December 06, 1954)
by Spahr, J Richard [53 pages; 1.7 MB] |
|
Abstract: No Abstract Available |
| 796. |
Free-flight determination of
force and stability characteristics of an inclined body of fineness
ratio 16.9 at a Mach number of 1.74 (November 15, 1954)
by Gillespie, Warren, Jr [18 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 797. |
Flight measurements of average
skin-friction coefficients on a parabolic body of revolution
(NACA-RM-10) at mach numbers from 1.0 to 3.7 (1954) by
Loposer, J. Dan, Rumsey, Charles B. [33 pages; 1.4 MB] |
|
Abstract: (abstract not available) |
| 798. |
Effects of subsonic Mach number
on the forces and pressure distributions on four NACA 64a-series
airfoil sections at angles of attack as high as 28 degrees
(1954) by Stiver, Louis S. Jr. [146 pages; 8.3 MB] |
|
Abstract: (abstract not available) |
| 799. |
Some observations of
shock-induced turbulent separation on supersonic diffusers
(1954) by Nussdorfer, Theodore J. [16 pages; 0.7 MB] |
|
Abstract: A survey of experimental data at
supersonic speed indicated that shock-induced separation of a turbulent
boundary layer will result for Mach numbers of approximately 1.33 or
greater when a theoretical stream static-pressure-rise ratio of
approximately 1.89 occurs across a shock interacting with the boundary
layer. The significance of this tentative criterion for turbulent
boundary-layer separation is discussed with respect to the design of
supersonic diffusers. |
| 800. |
Investigation of a three-blade
propeller in combination with two different spinners and an NACA D-type
cowl at Mach numbers up to 0.80 (1954) by Reynolds,
Robert M., Kenyon, George C. [64 pages; 6 MB] |
|
Abstract: (abstract not available) |
| 801. |
Flight Investigation of Engine
Nacelles and Wing Vertical Position on the Drag of a Delta-Wing
Airplane Configuration from Mach Number 0.8 to 2.0 (1954)
by Joseph H. Judd (Langley Aeronautical Laboratory, Langley Field, Va.)
[41 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 802. |
Theoretical prediction of
pressure distributions on nonlifting airfoils at high subsonic speeds
(1955) by John R. Spreiter, Alberta Alksne [44 pages; 4.1 MB] |
|
Abstract: Theoretical pressure distributions on
nonlifting circular-arc airfoils in two-dimensional flows with high
subsonic free-stream velocity are found by determining approximate
solutions, through an iteration process, of an integral equation for
transonic flow proposed by Oswatitsch. The integral equation stems
directly from the small-disturbance theory for transonic flow. This
method of analysis possesses the advantage of remaining in the
physical, rather than the hodograph, variable and can be applied in
airfoils having curved surfaces. After discussion of the derivation of
the integral equation and qualitative aspects of the solution, results
of calculations carried out for circular-arc airfoils in flows with
free-stream Mach numbers up to unity are described. These results
indicate most of the principal phenomena observed in experimental
studies. |
| 803. |
Measurement and analysis of
wing and tail buffeting loads on a fighter airplane (1955)
by Wilber B. Huston, T. H. Skopinski [28 pages; 2.8 MB] |
|
Abstract: The buffeting loads measured on the
wing and tail of a fighter airplane during 194 maneuvers are given in
tabular form, along with the associated flight conditions. Measurements
were made at altitudes of 30,000 to 10,000 feet and at speeds up to a
Mach number of 0.8. Least-squares methods have been used for a
preliminary analysis of the data. The agreement between the results of
this analysis and the loads measured in stalls is sufficiently good to
suggest the examination of the buffeting of other airplanes on the same
basis. |
| 804. |
A free-flight wind tunnel for
aerodynamic testing at hypersonic speeds (1955) by Alvin
Seiff [18 pages; 1.8 MB] |
|
Abstract: The supersonic free-flight wind tunnel
is a facility at the Ames Laboratory of the NACA in which aerodynamic
test models are gun-launched at high speed and directed upstream
through the test section of a supersonic wind tunnel. In this way, test
Mach numbers up to 10 have been attained and indications are that still
higher speeds will be realized. An advantage of this technique is that
the air and model temperatures simulate those of flight through the
atmosphere. Also the Reynolds numbers are high. Aerodynamic
measurements are made from photographic observation of the model
flight. Instruments and techniques have been developed for measuring
the following aerodynamic properties: drag, initial lift-curve slope,
initial pitching-moment-curve slope, center of pressure, skin friction,
boundary-layer transition, damping in roll, and aileron effectiveness. |
| 805. |
An investigation of the maximum
lift of wings at supersonic speeds (1955) by James J.
Gallagher, James N. Mueller [28 pages; 1.4 MB] |
|
Abstract: This report presents the results of an
exploratory investigation carried out in the Langley 9-inch supersonic
tunnel to determine the maximum lift of wings operating at supersonic
speeds. A variety of wing plan forms of random thickness distributions
were tested at Mach numbers of 1.55, 1.90, and 2.32 and at Reynolds
numbers varying between 0.74 x 10(6) and 0.27 x 10(6) at angles of
attack ranging from zero up through the angle at which maximum lift
occurred. Subsequent pressure-distribution tests on wings of triangular
and rectangular plan forms were made at a Mach number of 2.40. The
results of these tests substantiated the values of maximum lift
obtained during the force tests and further showed no appreciable
center-of-pressure shift over the entire angle-of-attack range. |
| 806. |
Generalized indical forces on
deforming rectangular wings in supersonic flight (1955)
by Harvard Lomax, Franklyn B. Fuller, Loma Sluder [28 pages; 2.4 MB] |
|
Abstract: A method is presented for determining
the time-dependent flow over a rectangular wing moving with a
supersonic forward speed and undergoing small vertical distortions
expressible as polynomials involving spanwise and chordwise distances.
The solution for the velocity potential is presented in a form
analogous to that for steady supersonic flow having the familiar
reflected area concept discovered by Evvard. Particular attention is
paid to indicial-type motions and results are expressed in terms of
generalized indicial forces. Numerical results for Mach numbers equal
to 1.1 and 1.2 are given for polynomials of the first and fifth degree
in the chordwise and spanwise directions, respectively, on a wing
having an aspect ratio of 4. |
| 807. |
Shock-turbulence interaction
and the generation of noise (1955) by H. S. Ribner [22
pages; 2 MB] |
|
Abstract: Interaction of convected field of
turbulence with shock wave is analyzed to yield modified turbulence,
entropy spottiness, and noise generated downstream of the shock.
Analysis is generalization of single-spectrum-wave treatment of
NACA-TN-2864. Formulas for spectra and correlations are obtained.
Numerical calculations yield curves of rms velocity components,
temperature, pressure, and noise in db against Mach number for m = 1 to
infinity; both isotropic and strongly axisymmetric
(lateral/longitudinal = 36/1) initial turbulence are treated. In either
case, turbulence of 0.1 percent longitudinal component generates about
120 dbs of noise. |
| 808. |
On the Kernel function of the
integral equation relating the lift and downwash distributions of
oscillating finite wings in subsonic flow (1955) by
Charles E. Watkins, Harry L. Runyan, Donald S. Woolston [16 pages; 1.3
MB] |
|
Abstract: This report treats the Kernel function
of an integral equation that relates a known prescribed downwash
distribution to an unknown lift distribution for a harmonically
oscillating finite wing in compressible subsonic flow. The Kernel
function is reduced to a form that can be accurately evaluated by
separating the Kernel function into two parts: a part in which the
singularities are isolated and analytically expressed and a nonsingular
part which may be tabulated. The form of the Kernel function for the
sonic case (Mach number 1) is treated separately. In addition, results
for the special cases of Mach number of 0 (incompressible case) and
frequency of 0 (steady case) are given. The derivation of the integral
equation which involves this Kernel function is reproduced as an
appendix. Another appendix gives the reduction of the form of the
Kernel function obtained herein for the three-dimensional case to a
known result of Possio for two-dimensional flow. A third appendix
contains some remarks on the evaluation of the Kernel function, and a
fourth appendix an alternate form of expression for the Kernel
function. |
| 809. |
Arrangement of fusiform bodies
to reduce the wave drag at supersonic speeds (1955) by
Morris D. Friedman, Doris Cohn [8 pages; 0.6 MB] |
|
Abstract: By means of linearized-body theory and
reverse-flow theorems, the wave drag of a system of fusiform bodies at
zero angle of attack and supersonic speeds is studied to determine the
effect of varying the relative location of the component parts. The
investigation is limited to two-body and three-body arrangements of
Sears-Haack minimum-drag bodies. It is found that in certain
arrangements the interference effects are beneficial, and may even
result in the two or three-body system having no more wave drag than
that of the principal body alone. The most favorable location appears
to be one in which the maximum cross-section of the auxiliary body is
slightly forward of the Mach cone from the tail of the main body. The
least favorable is the region between the Mach cone from the nose and
the forecone from the tail of the main body. |
| 810. |
Investigations at supersonic
speeds of 22 triangular wings representing two airfoil sections for
each of 11 apex angles (1955) by Eugene S. Love [60
pages; 3.1 MB] |
|
Abstract: The results of tests of 22 triangular
wings, representing two leading-edge shapes for each of 11 apex angles,
at Mach numbers 1.62, 1.92, and 1.40 are presented and compared with
theory. All wings have a common thickness ratio of 8 percent and a
common maximum-thickness point at 18 percent chord. Lift, drag, and
pitching moment are given for all wings at each Mach number. The
relation of transition in the boundary layer, shocks on the wing
surfaces, and characteristics of the pressure distributions is
discussed for several wings. |
| 811. |
An investigation of the effects
of heat transfer on boundary-layer transition on a parabolic body of
revolution (NACA RM-10) at a Mach number of 1.61 (1955)
by K. R. Czarnecki, Archibald R. Sinclair [12 pages; 1.2 MB] |
|
Abstract: Report presents the results of an
investigation conducted to determine the effects of heat transfer on
boundary-layer transition on a parabolic body of revolution (NACA rm-10
without fins) at Mach number of 1.61 and over a Reynolds number range
from 2.5 x 10(6) to 35 x 10(6). The maximum cooling of the model used
in these tests corresponded to a temperature ratio (ratio of
model-surface temperature to free-stream temperature) of 1.12, a value
somewhat higher than the theoretical value required for infinite
boundary-layer stability at this Mach number. The maximum heating
corresponded to a temperature ratio of about 1.85. Included in the
investigation was a study of the effects of surface irregularities and
disturbances generated in the airstream on the ability of heat transfer
to influence boundary-layer transition. |
| 812. |
Transonic flow past cone
cylinders (1955) by George E. Solomon [16 pages; 1.4 MB]
|
|
Abstract: Experimental results are presented for
transonic flow post cone-cylinder, axially symmetric bodies. The drag
coefficient and surface Mach number are studied as the free-stream Mach
number is varied and, wherever possible, the experimental results are
compared with theoretical predictions. Interferometric results for
several typical flow configurations are shown and an example of
shock-free supersonic-to-subsonic compression is experimentally
demonstrated. The theoretical problem of transonic flow past finite
cones is discussed briefly and an approximate solution of the axially
symmetric transonic equations, valid for a semi-infinite cone, is
presented. |
| 813. |
The dynamic-response
characteristics of a 35 degree swept-wing airplane as determined from
flight measurements (1955) by William C. Triplett,
Stuart C. Brown, G. Allan Smith [26 pages; 1.9 MB] |
|
Abstract: The longitudinal and
lateral-directional dynamic-response characteristics of a 35 degree
swept-wing fighter-type airplane determined from flight measurements
are presented and compared with predictions based on theoretical
studies and wind-tunnel data. Flights were made at an altitude of
35,000 feet covering the Mach number range of 0.50 to 1.04. A limited
amount of lateral-directional data were also obtained at 10,000 feet.
The flight consisted essentially of recording transient responses to
pilot-applied pulsed motions of each of the three primary control
surfaces. These transient data were converted into frequency-response
form by means of the Fourier transformation and compared with predicted
responses calculated from the basic equations. Experimentally
determined transfer functions were used for the evaluation of the
stability derivatives that have the greatest effect on the dynamic
response of the airplane. The values of these derivatives, in most
cases, agreed favorably with predictions over the Mach number range of
the test. |
| 814. |
A preliminary investigation of
aerodynamic characteristics of small inclined air outlets at transonic
Mach numbers (May 1955) by Paul E. Dewey [22 pages; 0.7
MB] |
|
Abstract: The aerodynamic characteristics of
several outlets with inclined or curved axes discharging air into a
transonic stream have been investigated. The data presented herein show
the discharge coefficient of such outlets and the static-pressure
distribution in the vicinity of the outlets for several values of
stream Mach number and discharge flow parameter. |
| 815. |
Application of the generalized
shock-expansion method to inclined bodies of revolution traveling at
high supersonic airspeeds (Apr 1955) by Raymond C. Savin
[72 pages; 2.1 MB] |
|
Abstract: The generalized shock-expansion method
is applied to obtain solutions to the flow field about pointed bodies
of revolution at high supersonic airspeeds and small angles of attack.
Simple explicit expressions are obtained for the surface Mach numbers
and surface pressures in the special case of slender bodies. In the
case of inclined cones, algebraic solutions are obtained defining the
entire flow field. Experimental pressure-distribution data for cones
and ogives at Mach numbers from 3 to 5 are included. |
| 816. |
Flight measurements of base
pressure on bodies of revolution with and without simulated rocket
chambers (Apr 1955) by Robert F. Peck [19 pages; 0.5 MB]
|
|
Abstract: Base pressures were measured on
fin-stabilized bodies of revolution with and without rocket chambers
and with and without a converging afterbody. At Mach numbers between
0.7 and 1.2, the results show that the presence of a cold rocket
chamber increased the pressure (less suction) over the center portion
of the bases. The effects of rocket chambers on pressures near the edge
of the bases were not as consistent throughout the Mach number range
nor as appreciable at most speeds as were the effects of pressures
measured on the center line. |
| 817. |
Turbulent-heat-transfer
measurements at a Mach number of 2.06 (Mar 1955) by
Maurice J. Brevoort, Bernard Rashis [21 pages; 0.5 MB] |
|
Abstract: Turbulent-heat-transfer measurements
were obtained through the use of an axially symmetric annular nozzle
which consists of an inner shaped center body and an outer cylindrical
sleeve. |
| 818. |
Pressure distributions on
triangular and rectangular wings to high angles of attack Mach numbers
2.46 and 3.36 (January 18, 1955) by Kaattari, George E
[31 pages; 1.3 MB] |
|
Abstract: Pressure distributions were measured
over rectangular wings of aspect ratio 2 and triangular wings of aspect
ratios 2 and 4 at Mach numbers of 2.46 and 3.36. The investigation
includes some comparison of the effects of Mach number, Reynolds
number, and thickening the wing root sections on the loading. |
| 819. |
The effect of a 4-percent-high
spoiler on buffeting forces on a naca 65(sub 06)A004 two-dimensional
airfoil at subsonic Mach numbers (March 23, 1955) by
Mellenthin, Jack A [15 pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 820. |
Experimental investigation of
some aerodynamic effects of a gap between wing and body of a moderately
slender wing-body combination at a Mach number of 1.4 (May 27,
1955) by Dugan, Duane W [35 pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 821. |
Investigation of some wake
vortex characteristics of an inclined ogive-cylinder body at Mach
number 1.98 (August 23, 1955) by Jorgensen, Leland H
Perkins, Edward W [48 pages; 1.8 MB] |
|
Abstract: No Abstract Available |
| 822. |
Drag and rolling-moment
effectiveness of trailing-edge spoilers at Mach numbers 2.2 and 5.0
(October 03, 1955) by Canning, Thomas N Derose, Charles E [51
pages; 1.6 MB] |
|
Abstract: No Abstract Available |
| 823. |
An investigation of the effects
of nose and lip shapes for an underslung scoop inlet at Mach numbers
from 0 to 1.9 (November 18, 1955) by Pfyl, Frank A [61
pages; 1.5 MB] |
|
Abstract: No Abstract Available |
| 824. |
Temperature recovery factors on
a slender 12 degree cone-cylinder at Mach numbers from 3.0 to 6.3 and
angles of attack up to 45 degrees (October 03, 1955) by
Reller, John O Hamaker, Frank M [57 pages; 1.9 MB] |
|
Abstract: No Abstract Available |
| 825. |
Effects of boundary-layer
separation on normal force and center of pressure of a cone-cylinder
model with a large base flare at Mach numbers from 3.00 to 6.28
(October 03, 1955) by Dennis, David H Syvertson, Clarence A [15
pages; 0.3 MB] |
|
Abstract: No Abstract Available |
| 826. |
An experimental investigation
of the hinge-moment characteristics of a constant-chord control surface
oscillating at high frequency (December 1955) by Reese,
David E JR Carlson, William C A [29 pages; 1.1 MB] |
|
Abstract: The results of an experimental
investigation of the hinge-moment characteristics of a constant-chord
control surface oscillating at high frequency is presented. The control
surface was mounted on an aspect-ratio-2 triangular wing. The
aerodynamic restoring-moment coefficient and damping-moment coefficient
were determined at a frequency of 260 cycles per second for a Mach
number range of 0.6 to 0.8 and 1.3 to 1.9 at angles of attack of 5
degrees and 10 degrees. The test results showed linear theory to be a
reliable guide to the prediction of the trend of the restoring-moment
coefficient with Mach number for the supersonic speed range of the
investigation but overestimated the magnitude of the coefficient. The
experimental values of the damping-moment coefficient were, for the
most part, more positive than those indicated by the theory and, for
some conditions, could lead to instability of the control surface.
Comparison of the results of this investigation with those of previous
investigations at 0 and 50 cycles per second showed that frequency had
little effect on the restoring-moment coefficient. The damping-moment
coefficient was similarly insensitive to frequency at an oscillation
amplitude of plus-or-minus 1.0 degrees but at an amplitude
ofplus-or-minus 2.5 degrees the results showed a destabilizing shift
with increasing frequency. |
| 827. |
Experimental investigation of
methods of improving diffuser-exit total-pressure profiles for
side-inlet model at Mach number 3.05 (August 29, 1955)
by Piercy, Thomas G Klann, John L [43 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 828. |
Free-flight heat-transfer
measurements on two 20 degree-cone-cylinders at Mach numbers from 1.3
to 4.9 (July 18, 1955) by Rabb, Leonard Simpkinson,
Scott H [59 pages; 4.9 MB] |
|
Abstract: No Abstract Available |
| 829. |
Performance characteristics of
axisymmetric two-cone and isentropic nose inlets at Mach number 1.90
(December 07, 1955) by Conners, James F Meyer, Rudolph C [33
pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 830. |
Preliminary investigation of
some internal boundary-layer-control systems on a side inlet at Mach
number 2.96 (February 18, 1955) by Piercy, Thomas G [38
pages; 2 MB] |
|
Abstract: No Abstract Available |
| 831. |
Wind-tunnel investigation at
Mach 1.9 of multijet-missile base pressures (March 1955)
by Baughman, L Eugene [14 pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 832. |
Investigation of a ramp-type
inlet designed for improved angle-of-attack performance at Mach number
2.0 (February 23, 1955) by Wise, G A Campbell, R C [15
pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 833. |
Application of oblique-shock
sensing system to ram-jet-engine flight Mach number control
(March 03, 1955) by Wilcox, Fred A Hearth, Donald P [30 pages;
1.1 MB] |
|
Abstract: No Abstract Available |
| 834. |
Effect of centerbody
boundary-layer removal near the throat of three coniccal nose inlets at
Mach 1.6 to 2.0 (November 15, 1955) by Kremzier, Emil J
Wise, George A [19 pages; 0.5 MB] |
|
Abstract: A zero angle-of-attack investigation of
the effect of compression-surface boundary-layer bleed through
perforations near the throat of three full-scale conical nose inlets
was conducted in the Lewis 8- by 6- foot supersonic wind tunnel for a
Mach number range from 1.6 to 2.0. The bleed system increased pressure
recovery, shifted the peak of the diffuser-discharge total-pressure
profile toward the center-body, and decreased the range of stable inlet
operation. A propulsion-system thrust minus drag analysis indicated
that the increases in inlet pressure recovery were too small to
compensate for the esimated bleed system drags. |
| 835. |
Boundary-layer transition at
high Reynolds numbers as obtained in flight of a 20 degree
cone-cylinder with wall to local stream temperature ratios near 1.0
(November 03, 1955) by Rabb, Leaonard Disher, John H [36 pages; 1
MB] |
|
Abstract: Boundary-layer transition data at low
ratios of wall to local stream temperature have been obtained during
the free flight of a highly polished cone-cylinder to a maximum Mach
number of 5.02 A maximum transition Reynolds number of 32 x 10(exp 6)
occurred at a distance of 25.84 inches from the cone apex. The
temperature ratio at transition for a local Mach number of 4.0 was
approximately 1.30 as compared with theoretical infinite stability
solutions of 1.47 and 1.65 by Dunn and Lin (three-dimensional) and Van
Driest (two-dimensional), respectively. |
| 836. |
Preliminary investigation of
effect on performance of dividing conical-spike nose inlets into halves
at Mach numbers 1.5 to 2.0 (December 19, 1955) by Allen,
John L [21 pages; 0.6 MB] |
|
Abstract: Inserting a splitter plate in the
subsonic diffuser caused a pressure-recovery loss of about 1 percent
for an inlet with a long nearly constant-area throat section. The loss
was due to the increased surface area. Another inlet, which had a
comparatively rapid area increase immediately after the throat,
experienced pressure-recovery losses of 5 and 6 percent at Mach numbers
of 1.8 and 2.0, respectively, and about 1 percent at Mach 1.5. |
| 837. |
Jet effects on longitudinal
trim of an airplane configuration measured at Mach numbers between 1.2
and 1.8 (January 18, 1955) by Peck, Robert F [18 pages;
0.5 MB] |
|
Abstract: No Abstract Available |
| 838. |
Free-flight investigation,
including some effects of wing aeroelasticity, of the rolling
effectiveness of an all-movable horizontal tail with differential
incidence at Mach numbers from 0.6 to 1.5 (January 25, 1955)
by English, Roland D [12 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 839. |
Turbulent convective
heat-transfer coefficients measured from flight tests of four research
models (NACA RM-10) at Mach numbers from 1.0 to 3.6 (March 11,
1955) by Chauvin, Leo T Maloney, Joseph P [31 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 840. |
Performance measurements from a
rocket-powered exploratory research missile flown to a Mach number of
10.4 (March 15, 1955) by Piland, Robert O [13 pages; 0.5
MB] |
|
Abstract: No Abstract Available |
| 841. |
Flutter experiences with thin
pointed-tip wings during flight tests of rocket-propelled models at
Mach numbers from 0.8 to 1.95 (April 04, 1955) by
Wallskog, Harvey A [33 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 842. |
Aerodynamic-heating data
obtained from free-flight tests between Mach numbers of 1 and 5
(March 11, 1955) by Rumsey, Charles B Piland, Robert O Hopko,
Russell N [22 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 843. |
Aerodynamic characteristics of
a 60 degree delta wing having a half-delta tip control at a Mach number
of 4.04 (April 25, 1955) by Ulmann, Edward F Smith, Fred
M [27 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 844. |
Flight and preflight tests of a
ram jet burning magnesium slurry fuel and utilizing a solid-propellant
gas generator for fuel expulsion (April 06, 1955) by
Bartlett, Walter, A , jr Hagginbotham, William K , Jr [40 pages; 0.9
MB] |
|
Abstract: Data obtained from the first flight
test of a ram jet utilizing a magnesium slurry fuel are presented. The
ram jet accelerated from a Mach number of 1.75 to a Mach number of 3.48
in 15.5 seconds. During this period a maximum values of air specific
impulse and gross thrust coefficient were calculated to be 151 seconds
and 0.658, respectively. The rocket gas generator used as a
fuel-pumping system operated successfully. |
| 845. |
Preliminary results of an
investigation at transonic speeds to determine the effects of a heated
propulsive jet on the drag characteristics of a related series of
afterbodies (March 25, 1955) by Henry, Beverly Z , Jr
Cahn, Maurice S [29 pages; 2 MB] |
|
Abstract: Preliminary results are presented from
an investigation to determine the influence of afterbody geometry on
the effects of a sonic propulsive jet at transonic speeds. The results
presented are base pressure coefficient and afterbody pressure-drag
coefficient as a function of jet pressure ratio for different values of
Mach number and jet temperature. Geometric parameters investigated
include boattail angle, jet-to-model diameter ratio, and jet-to-base
diameter ratio. |
| 846. |
An investigation of the
aerodynamic characteristics of thin delta wings with a symmetrical
double-wedge section at a Mach number of 6.9 (October 14,
1955) by Bertram, Mitchel H Mccauley, William D [41 pages; 1.4
MB] |
|
Abstract: No Abstract Available |
| 847. |
Investigation of interference
lift, drag, and pitching moment of a series of triangular wing and body
combinations at a Mach number of 1.62 (May 27, 1955) by
Coletti, Donard E [51 pages; 1.6 MB] |
|
Abstract: No Abstract Available |
| 848. |
An evaluation of a
rolleron-roll-rate-stabilization system for a canard missile
configuration at Mach numbers from 0.9 to 2.3 (September 15,
1955) by Nason, Martin L Rock, Rupert S Brown, Clarence, Jr [48
pages; 1.8 MB] |
|
Abstract: This type of damper provides roll
damping by the action of gyro-actuated uncoupled wing-tip ailerons. A
dynamic roll instability predicted by the analysis was confirmed by
flight testing and was subsequently eliminated by introducing
control-surface damping about the rolleron hinge line. |
| 849. |
Experimental drag coefficients
of round noses with conical windshields at Mach number 2.72
(June 28, 1955) by Jones, Jim J [19 pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 850. |
An experimental investigation
at a Mach number of 2.01 of the effects of body cross-section shape on
the aerodynamic characteristics of bodies and wing-body combinations
(July 21, 1955) by Carlson, Harry W Gapcynski, John P [30 pages;
2.1 MB] |
|
Abstract: No Abstract Available |
| 851. |
Investigation of equilibrium
temperatures and average laminar heat-transfer coefficients for the
front half of swept circular cylinders at a Mach number of 6.9
(August 18, 1955) by Feller, William V [22 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 852. |
A free-flight investigation of
the effects of simulated sonic turbojet exhaust on the drag of a
boattail body with various jet sizes from Mach number 0.87 to 1.50
(August 18, 1955) by Falanga, Ralph A [24 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 853. |
Tests of aerodynamically heated
multiweb wing structures in a free jet at Mach number 2 : two
aluminum-alloy models of 20-inch chord with 0.064- and 0.081-inch-thick
skin (August 09, 1955) by Griffith, George E
Miltonberger, Georgene H Rosecrans, Richard [40 pages; 1.3 MB] |
|
Abstract: No Abstract Available |
| 854. |
Collection and summary of
flap-type-aileron rolling-effectiveness data at zero lift as determined
by rocket-powered model tests at Mach numbers between 0.6 and 1.6
(September 02, 1955) by Strass, H Kurt Stephens, Emily W Fields,
E M Schult, Eugene D [96 pages; 3.2 MB] |
|
Abstract: No Abstract Available |
| 855. |
Free-flight measurements of
aerodynamic heat transfer to Mach number 3.9 and of drag to Mach number
6.9 of a fin-stabilized cone-cylinder configuration (October
07, 1955) by Rumsey, Charles B [22 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 856. |
Flight investigation at
supersonic Mach numbers of an automatic acceleration control missile in
which rate damping is obtained from a linear accelerometer placed ahead
of the missile center of gravity (November 08, 1955) by
Seaberg, Ernest C Sproull, Royce H Reid, H J E , Jr [35 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 857. |
Free-flight tests to determine
the power-on and power-off pressure distribution and drag of the NACA
RM-10 research vehicle at large Reynolds numbers between Mach numbers
0.8 and 3.0 (September 20, 1955) by Hoffman, Sherwood
[56 pages; 4 MB] |
|
Abstract: No Abstract Available |
| 858. |
Aerodynamic loads on an
external store adjacent to a 45 degree sweptback wing at Mach numbers
from 0.70 to 1.96, including an evaluation of techniques used
(November 15, 1955) by Guy, Lawrence D Hadaway, William M [110
pages; 2.7 MB] |
|
Abstract: Aerodynamic forces and moments have
been obtained in the Langley 9- by 12-inch blowdown tunnel on an
external store and on a 45 degree swept-back wing-body combination
measured separately at Mach numbers from 0.70 to 1.96. The wing was
cantilevered and had an aspect ratio of 4.0; the store was
independently sting-mounted and had a Douglas Aircraft Co. (DAC) store
shape. The angle of attack range was from -3 degrees to 12 degrees and
the Reynolds number (based on wing mean aerodynamic chord) varied from
1.2 x10(6) to 1.7 x 10(6). Wing-body transonic forces and moments have
been compared with data of a geometrically similar full-scale model
tested in the Langley 16-foot and 8-foot transonic tunnels in order to
aid in the evaluation of transonic-tunnel interference. The principal
effect of the store, for the position tested, was that of delaying the
wing-fuselage pitch-up tendency to higher angles of attack at Mach
numbers from 0.70 to 0.90 in a manner similar to that of a wing chord
extension. The most critical loading condition on the store was that
due to side force, not only because the loads were of large magnitude
but also because they were in the direction of least structural
strength of the supporting pylon. These side loads were greatest at
high angles of attack in the supersonic speed range. Removal of the
supporting pylon (or increasing the gap between the store and wing)
reduced the values of the variation of side-force coefficientwith angle
of attack by about 50 percent at all test Mach numbers, indicating that
important reductions in store side force may be realized by proper
design or location of the necessary supporting pylon. A change of the
store skew angle (nose inboard) was found to relieve the excessive
store side loads throughout the Mach number range. It was also
determined that the relative position of the fuselage nose to the store
can appreciably affect the store side forces at supersonic speeds. |
| 859. |
An experimental investigation
of the flow phenomena over bodies at high angles of attack at a Mach
number of 2.01 (October 27, 1955) by Gapcynski, John P
[25 pages; 0.7 MB] |
|
Abstract: No Abstract Available |
| 860. |
Lift, drag, and longitudinal
stability at Mach numbers from 1.4 to 2.3 of a rocket-powered model
having a 52.5 degree sweptback wing of aspect ratio 3 and inline tail
surfaces (December 15, 1955) by Gillespie, Warren, Jr
[31 pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 861. |
Flight investigation of the
effect of underwing propulsive jets on the lift, drag, and longitudinal
stability of a delta-wing configuration at Mach numbers from 1.23 to
1.62 (December 15, 1955) by Falanga, Ralph A Judd,
Joseph H [33 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 862. |
Investigation of interference
lift, drag, and pitching moment of a series of triangular-wing and body
combinations at a Mach number of 1.94 (December 21, 1955)
by Coletti, Donald E [53 pages; 1.7 MB] |
|
Abstract: No Abstract Available |
| 863. |
An experimental investigation
of the base pressure characteristics of nonlifting bodies of revolution
at Mach numbers from 2.73 to 4.98 (March 17, 1955) by
Reller, John O , Jr Hamaker, Frank M [46 pages; 1.5 MB] |
|
Abstract: No Abstract Available |
| 864. |
An investigation of several
NACA 1-series nose inlets with and without protruding central bodies at
high-subsonic Mach numbers and at a Mach number of 1.2 (May
1955) by Pendley, Robert E Robinson, Harold L [52 pages; 1.5 MB] |
|
Abstract: No Abstract Available |
| 865. |
An investigation of string
support interference on base pressure and forebody chord force at Mach
numbers from 0.60 to 1.30 (January 28, 1955) by Tunnell,
Phillips J [20 pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 866. |
The longitudinal
characteristics at Mach numbers up to 0.92 of several
wing-fuselage-tail combinations having sweptback wings with NACA
four-digit thickness distributions (March 24, 1955) by
Sutton, Fred B Dickson, Jerald K [129 pages; 3.5 MB] |
|
Abstract: No Abstract Available |
| 867. |
Effects of sweep and taper on
the longitudinal characteristics of an aspect ratio 3 wing-body
combination at Mach numbers from 0.6 to 1.4 (March 23, 1955)
by Knechtel, Earl D Summers, James L [38 pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 868. |
Downwash survey behind two
low-aspect-ratio variable-incidence wings in combination with three
different size fuselages at a Mach number of 0.25 (March 30,
1955) by Hopkins, Edward J Sorensen, Norman E [55 pages; 2.8 MB] |
|
Abstract: No Abstract Available |
| 869. |
The unsteady normal-force
characteristics of selected NACA profiles at high subsonic Mach numbers
(May 27, 1955) by Polentz, Perry P Page, William A
Levy, Lionel L , Jr [111 pages; 3.6 MB] |
|
Abstract: No Abstract Available |
| 870. |
A correlation of airfoil
section data with the aerodynamic loads measured on a 45 degree
sweptback wing model at subsonic Mach numbers (May 27, 1955)
by Walker, Harold J Maillard, William C [82 pages; 2.6 MB] |
|
Abstract: No Abstract Available |
| 871. |
Lift, drag, and static
longitudinal stability characteristics of four airplane-like
configurations at Mach numbers from 3.00 to 6.28 (April 25,
1955) by Neice, Stanford E Wong, Thomas J Hermach, Charles A [19
pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 872. |
The effect of wing fences on
the longitudinal characteristics at Mach numbers up to 0.92 of a
wing-fuselage-tail combination having a 40 degree sweptback wing with
NACA 64A thickness distribution (May 27, 1955) by
Dickson, Gerald K Sutton, Fred B [55 pages; 1.5 MB] |
|
Abstract: No Abstract Available |
| 873. |
Experimental investigation at
Mach numbers from 0 to 1.9 of trapezoidal and circular side inlets for
a fighter-type airplane (July 28, 1955) by Mossman,
Emmett A Pfyl, Frank A Lazzeroni, Frank A [40 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 874. |
A comparison at Mach numbers up
to 0.92 of the calculated and experimental downwash and wake
characteristics at various horizontal tail heights behind a wing with
45 degree of sweepback (June 28, 1955) by Stephenson,
Jack D Selan, Ralph Bandettini, Angelo [83 pages; 2.3 MB] |
|
Abstract: No Abstract Available |
| 875. |
Static stability and control of
canard configurations at Mach numbers from 0.70 to 2.22 :
lateral-directional characteristics of a triangular wing and canard
(March 28, 1955) by Peterson, Victor L Menees, Gene P [78 pages;
2.9 MB] |
|
Abstract: No Abstract Available |
| 876. |
Lift, drag, and longitudinal
stability at Mach numbers from 0.8 to 2.1 of a rocket-powered model
having a tapered unswept wing of aspect ratio 3 and inline tail
surfaces (April 25, 1955) by Gillespie, Warren, Jr [30
pages; 1 MB] |
|
Abstract: No Abstract Available |
| 877. |
Effect of convergent ejector
nozzles on the boattail drag of a 16 degree conical afterbody at Mach
numbers of 0.6 to 1.26 (September 17, 1955) by Cubbage,
James M , Jr [35 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 878. |
Optimum flight paths of
turbojet aircraft (1955) by Miele, Angelo. [48 pages;
1.1 MB] |
|
Abstract: The climb of turbojet aircraft is
analyzed and discussed including the accelerations. Three particular
flight performances are examined: minimum time of climb, climb with
minimum fuel consumption, and steepest climb. The theoretical results
obtained from a previous study are put in a form that is suitable for
application on the following simplifying assumptions: the Mach number
is considered an independent variable instead of the velocity; the
variations of the airplane mass due to fuel consumption are
disregarded; the airplane polar is assumed to be parabolic; the path
curvatures and the squares of the path angles are disregarded in the
projection of the equation of motion on the normal to the path; lastly,
an ideal turbojet with performance independent of the velocity is
involved. The optimum Mach number for each flight condition is obtained
from the solution of a sixth order equation in which the coefficients
are functions of two fundamental parameters: the ratio of minimum drag
in level flight to the thrust and the Mach number which represents the
flight at constant altitude and maximum lift-drag ratio. |
| 879. |
Preliminary experimental
investigation of a variable-area, variable-internal-contraction air
inlet at Mach numbers between 1.42 and 2.44 (1955) by
Scherrer, Richard., Gowen, Forrest E. [27 pages; 1 MB] |
|
Abstract: The performance of a rectangular cross
section, variable-area, variable-internal-contraction air inlet has
been investigated at zero angle of attack at Mach no.s of 1.42, 1.75,
1.90, 1.99 and 2.44. |
| 880. |
Comparison between analytical
and wind-tunnel results on flutter of several low-aspect-ratio,
high-density, unswept wings at high subsonic speeds and zero angle of
attack (1955) by Warner Robert W. [26 pages; 1.2 MB] |
|
Abstract: Experimental flutter Mach numbers have
been estimated for several unswept, cantilever wings from the results
of previous tests at zero angle of attack. |
| 881. |
Aerodynamic characteristics of
two rectangular-plan-form, all moveable controls in combination with a
slender body of revolution at Mach numbers from 3.00 to 6.25
(1955) by Wong, Thomas J., Gloria, Hermilo R. [40 pages; 1.8 MB] |
|
Abstract: Results of tests to determine the
aerodynamic characteristics of all-movable control and body
combinations at angles of attack from 0 degrees to 25 degrees and
control deflection angles from -30 degrees to +30 degrees are presented
and compared with theory. |
| 882. |
A study of local-pressure
fluctuations relative to static-pressure distributions on
two-dimensional airfoils at high subsonic Mach numbers (1955)
by Coe, Charles F. [68 pages; 2.2 MB] |
|
Abstract: The relationship of local-pressure
fluctuations to the time-average static-pressure distribution has been
investigated for six symetrical two-dimensional airfoils. |
| 883. |
Longitudinal Stability
Characteristics at Mach Numbers up to 0.92 of a Wing-Body-Tail
Combination Having a Wing with 45o of Sweepback and a Tail in Various
Vertical Positions (1955) by Jack D. Stephenson, Angelo
Bandettini, and Ralph Selan (Ames Aeroanutical Laboratory, Moffett
Field Calif.) [65 pages; 2.5 MB] |
|
Abstract: No Abstract Available |
| 884. |
Theoretical and experimental
investigation of the effect of tunnel walls on the forces on an
oscillating airfoil in two-dimensional subsonic compressible flow
(Jan 1956) by Harry L. Runyan, Donald S. Woolston, Gerald A.
Rainey [22 pages; 1.1 MB] |
|
Abstract: This report presents a theoretical and
experimental investigation of the effect of wind- tunnel walls on the
air forces on an oscillating wing in two-dimensional subsonic
compressible flow. A method of solving an integral equation which
relates the downwash on a wing to the unknown loading is given, and
some comparisons are made between the theoretical results and the
experimental results. A resonance condition, which was predicted by
theory in a previous report (NACA report 1150), is shown experimentally
to exist. In addition, application of the analysis is made to a number
of examples in order to illustrate the influence of walls due to
variations in frequency of oscillation, Mach number , and ratio of
tunnel height to wing semichord. |
| 885. |
A theory for stability and buzz
pulsation amplitude in ram jets and an experimental investigation
including scale effects (Jan 1956) by Robert L. Trimpi
[24 pages; 2.3 MB] |
|
Abstract: From a theory developed on a
quasi-one-dimensional-flow basis, it is found that the stability of the
ram jet is dependent upon the instantaneous values of mass flow and
total pressure recovery of the supersonic diffuser and immediate
neighboring subsonic diffuser. Conditions for stable and unstable flow
are presented. The theory developed in the report is in agreement with
the experimental data of NACA-TN-3506 and NACA-RM-L50K30. A simple
theory for predicting the approximate amplitude of small pressure
pulsation in terms of mass-flow decrement from minimum-stable mass flow
is developed and found to agree with experiments. Cold-flow tests at a
Mach number of 1.94 of ram-jet models having scale factors of 3.15:1
and Reynolds number ratios of 4.75:1 with several supersonic diffuser
configurations showed only small variations in performance between
geometrically similar models. The predominant variation in steady-flow
performance resulted from the larger boundary layer in the combustion
chamber of the low Reynolds number models. The conditions at which buzz
originated were nearly the same for the same supersonic diffuser
(cowling-position angle) configurations in both large and small
diameter models. There was no appreciable variation in stability limits
of any of the models when the combustion- chamber length was increased
by a factor of three. The unsteady-flow performance and wave patterns
were also similar when considered on a reduced-frequency basis
determined from the relative lengths of the model. The negligible
effect of Reynolds number on stability of the off-design configurations
was not anticipated in view of the importance of boundary layer to
stability, and this result should not be construed to be generally
applicable. (author) |
| 886. |
Theoretical investigation of
flutter of two-dimensional flat panels with one surface exposed to
supersonic potential flow (Jan 1956) by Herbert C.
Nelson, Herbert J. Cunningham [24 pages; 1.7 MB] |
|
Abstract: A Rayleigh type analysis involving
chosen modes of the panel as degrees of freedom is used to treat the
flutter of a two-dimensional flat panel supported at its leading and
trailing edges and subjected to a middle-plane tensile force. The panel
has a supersonic stream passing over its upper surface and still air
below. The aerodynamic forces due to the supersonic stream are obtained
from the theory for linearized two-dimensional unsteady flow and the
forces due to the still air are obtained from acoustical theory. In
order to study the effect of increasing the number of modes in the
analysis, two and then four modes are employed. The modes used are the
first four natural modes of the panel in a vacuum with no tensile force
acting. The analysis includes these variables: Mach number, structural
damping, tensile force, density of the still air, and edge fixity
(clamped and pinned). For certain combinations of these variables,
stability boundaries are obtained which can be used to determine the
panel thickness required to prevent flutter for any panel material and
altitude. |
| 887. |
Flight determination of drag of
normal-shock nose inlets with various cowling profiles at Mach numbers
from 0.9 to 1.5 (Jan 1956) by R. I. Sears, C. F. Merlet,
L. W. Putland [20 pages; 0.7 MB] |
|
Abstract: External-drag data are presented for
normal-shock nose inlets with NACA 1-series, parabolic, and conic
cowling profiles. The tests were made at an angle of attack of 0
degrees by using rocket-propelled models in free flight at Mach numbers
from 0.9 to 1.5. The Reynolds number based on body maximum diameter
varied from 2.5 x 10 sup 6 to 5.5 x 10 sup 6. At maximum flow rate, the
inlet models had about the same external drag at a Mach number of
approximately 1.1, but at higher Mach numbers the sharp-lip conic
cowling had the least drag. Blunting or beveling the lip of the conic
cowling while keeping the fineness ratio constant resulted in drag
coefficients slightly higher than for the sharp-lip conic cowling at
maximum flow rate. At a mass-flow ratio of about 0.8, the conic
cowlings with sharp, blunt, or beveled lips and the parabolic cowling
all gave about the same drag. The higher drag of the NACA 1-49-300
cowling, compared with the blunt-lip conic cowling, is associated with
the greater fullness back of the inlet. |
| 888. |
A special method for finding
body distortions that reduce the wave drag of wing and body
combinations at supersonic speeds (Jan 1956) by Harvard
Lomax, Max A. Heaslet [38 pages; 2.2 MB] |
|
Abstract: For a given wing and supersonic Mach
number, the problem of shaping an adjoining fuselage so that the
combination will have a low wave drag is considered. Only fuselages
that can be simulated by singularities (multipoles) distributed along
the body axis are studied. However, the optimum variations of such
singularities are completely specified in terms of the given wing
geometry. An application is made to an elliptic wing having a biconvex
section, a thickness-chord ratio equal to 0.05 at the root, and an
aspect ratio equal to 3. A comparison of the theoretical results with a
wind-tunnel experiment is also presented. |
| 889. |
Theory of wing-body drag at
supersonic speeds (Jan 1956) by Robert T. Jones [7
pages; 0.5 MB] |
|
Abstract: The relation of Whitcomb's (area rule)
to the linear formulas for wave drag at lightly supersonic speeds is
discussed. By adopting an approximate relation between the source
strength and the geometry of a wing-body combination, the wave-drag
theory is expressed in terms involving the areas intercepted by oblique
planes or Mach planes. The resulting formulas are checked by comparison
with the drag measurements obtained in wind-tunnel experiments and in
experiments with falling models in free air. Finally, a theory for
determining wing-body shapes of minimum drag at supersonic Mach numbers
is discussed and some preliminary experiments are reported. |
| 890. |
Intensity, scale, and spectra
of turbulence in mixing region of free subsonic jet (Jan 1956)
by James C. Laurence [28 pages; 2 MB] |
|
Abstract: Report presents the results of the
measurements of intensity of turbulence, the longitudinal and lateral
correlation coefficients, and the spectra of turbulence in a 3.5-inch-
diameter free jet measured with hot-wire anemometers at exit Mach
numbers from 0.2 to 0.7 and Reynolds numbers from 192,000 to 725,000. |
| 891. |
Similar solutions for the
compressible laminar boundary layer with heat transfer and pressure
gradient (Jan 1956) by Clarence B. Cohen, Eli Reshotko
[38 pages; 2.5 MB] |
|
Abstract: Stewartson's transformation is applied
to the laminar compressible boundary-layer equations and the
requirement of similarity is introduced, resulting in a set of ordinary
nonlinear differential equations previously quoted by Stewartson, but
unsolved. The requirements of the system are Prandtl number of 1.0,
linear viscosity-temperature relation across the boundary layer, an
isothermal surface, and the particular distributions of free-stream
velocity consistent with similar solutions. This system admits axial
pressure gradients of arbitrary magnitude, heat flux normal to the
surface, and arbitrary Mach numbers. The system of differential
equations is transformed to integral system, with the velocity ratio as
the independent variable. For this system, solutions are found by
digital computation for pressure gradients varying from that causing
separation to the infinitely favorable gradient and for wall
temperatures from absolute zero to twice the free-stream stagnation
temperature. Some solutions for separated flows are also presented. |
| 892. |
A factor affecting transonic
leading-edge flow separation (Oct 1956) by George P.
Wood, Paul B. Gooderum [44 pages; 1.2 MB] |
|
Abstract: A change in flow pattern that was
observed as the free-stream Mach number was increased in the vicinity
of 0.8 was described in NACA Technical Note 1211 by Lindsey, Daley, and
Humphreys. The flow on the upper surface behind the leading edge of an
airfoil at an angle of attack changed abruptly from detached flow with
an extensive region of separation to attached supersonic flow
terminated by a shock wave. In the present paper, the consequences of
shock-wave - boundary layer interaction are proposed as a factor that
may be important in determining the conditions under which the change
in flow pattern occurs. Some experimental evidence in support of the
importance of this factor is presented. |
| 893. |
Wind-tunnel calibration of a
combined pitot-static tube and vane-type flow-angularity indicator at
Mach numbers of 1.61 and 2.01 (Oct 1956) by Archibald R.
Sinclair, William D. Mace [12 pages; 0.4 MB] |
|
Abstract: A limited calibration of a combined
pitot-static tube and vane-type flow-angularity indicator has been made
in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers
of 1.61 and 2.01. The results indicated that the angle-of-yaw
indications were affected by unsymmetric shock effects at low angles of
attack. |
| 894. |
Charts adapted from Van
Driest's turbulent flat-plate theory for determining values of
turbulent aerodynamic friction and heat-transfer coefficients
(Oct 1956) by Dorothy B. Lee, Maxime A. Faget [17 pages; 0.8 MB] |
|
Abstract: A modified method of Van Driest's
flat-plate theory for turbulent boundary layer has been found to
simplify the calculation of local skin-friction coefficients which, in
turn, have made it possible to obtain through Reynolds analogy
theoretical turbulent heat-transfer coefficients in the form of Stanton
number. A general formula is given and charts are presented from which
the modified method can be solved for Mach numbers 1.0 to 12.0,
temperature ratios 0.2 to 6.0, and Reynolds numbers 0.2 times 10 to the
6th power to 200 times 10 to the 6th power. |
| 895. |
On slender-body theory and the
area rule at transonic speeds (Nov 1956) by Keith C.
Harder, E. B. Klunker [15 pages; 0.5 MB] |
|
Abstract: The basic ideas of the slender-body
approximation have been applied to the nonlinear transonic-flow
equation for the velocity potential in order to obtain some of the
essential features of slender-body theory at transonic speeds. The
results of the investigation are presented from a unified point of view
which demonstrates the similarity of slender-body solutions in the
various Mach number ranges. The transonic area rule and some conditions
concerning its validity follow from the analysis. |
| 896. |
Wind-tunnel investigation to
determine the horizontal- and vertical-tail contributions to the static
lateral stability characteristics of a complete-model swept-wing
configuration at high subsonic speeds (Nov 1956) by
James W. Wiggins, Richard E. Kahn, Paul G. Fournier [35 pages; 0.9 MB] |
|
Abstract: An investigation was conducted in the
Langley high-speed 7- by 10-foot tunnel to determine the horizontal-
and vertical-tail contributions to the static lateral stability of a
complete-model swept-wing configuration at high subsonic speeds. The
results indicate that, in a general, Mach number effects within the
range studied and wing effects on the tail contribution were small and
the overall trends of the data of the present investigation agreed with
those which have been established at low speeds. |
| 897. |
Flight techniques for
determining airplane drag at high Mach numbers (Aug 1956)
by De E. Beeler, Donald R. Bellman, Edwin J. Saltzman [41 pages; 1.3
MB] |
|
Abstract: The measurement of total airplane drag
in flight is necessary to assess the applicability of wind-tunnel model
data. The NACA High-Speed Flight Station has investigated and developed
techniques for measuring the drag of high-speed research airplanes and
current fighter-type airplanes. |
| 898. |
Some observations on maximum
pressure rise across shocks without boundary-layer separation on
airfoils at transonic speeds (Nov 1956) by Walter F.
Lindsey, Patrick J. Johnston [28 pages; 0.8 MB] |
|
Abstract: An investigation of the two-dimensional
flow along flat plates having rounded leading edges has provided
additional information on shock-induced separation. The results
indicate that laminar boundary layers can sustain the theoretical
pressure rise for normal shocks without separating provided that the
local Mach numbers are less than about 1.4. |
| 899. |
Some measurements of
aerodynamic forces and moments at subsonic speeds on a wing-tank
configuration oscillating in pitch about the wing midchord
(Dec 1956) by Sherman A. Clevenson, Sumner A. Leadbetter [39
pages; 1 MB] |
|
Abstract: Measurements are presented of the
aerodynamic forces and moments acting on a wing-tank configuration,
with or without fins, oscillating in pitch about the wing-root
midchord. The reduced-frequency range was from 0.050 to 0.657, whereas
the Mach number and Reynolds number ranges were from 0.18 to 0.75 and
0.9 X 10(6) to 9.5 10((6), respectively. |
| 900. |
Conversion of inviscid
normal-force coefficients in helium to equivalent coefficients in air
for simple shapes at hypersonic speeds (Oct 1956) by
James N. Mueller [32 pages; 0.9 MB] |
|
Abstract: A correlation factor applicable for
converting inviscid aerodynamic normal-force coefficients of simple
shapes in helium to equivalent coefficients in air is found by using
calculations based on the shock-expansion method at Mach numbers of 12,
16, and 20. |
| 901. |
Spreading characteristics of a
jet expanding from choked nozzles at mach 1.91 (Dec 1956)
by Morris D. Rousso, Eugene L. Baughman [28 pages; 0.9 MB] |
|
Abstract: Total-temperature surveys were made to
determine the gross spreading characteristics of jets expanding from
axisymmetric convergent and convergent-divergent nozzles in a
supersonic stream. The nozzles were installed in the base of conically
boattailed bodies of revolution. Surveys were made in a region between
the nozzle exit and a station 8 nozzle diameters downstream of the exit
for jet pressure ratios from 2.5 to 16.0. |
| 902. |
Effects of two trailing-edge
controls on the aerodynamic characteristics of a rectangular wing and
body combination at Mach numbers from 3.00 and 5.05 (February
1956) by Gloria, Hermilo R Wong, Thomas J [23 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 903. |
Lift, drag, and static
longitudinal stability characteristics of configurations consisting of
three triangular wing panels and a body of equal length at Mach numbers
from 3.00 to 6.28 (February 07, 1956) by Savin, Raymond
C Wong, Thomas J [20 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 904. |
External-store drag reduction
at transonic and low supersonic Mach numbers by application of
Baldwin's moment-of-area rule (March 05, 1956) by Levy,
Lionel L JR Dickey, Robert R [13 pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 905. |
A preliminary investigation of
the static stability characteristics of four airplane-like
configurations at Mach numbers from 3.00 to 6.28 (March 26,
1956) by Wong, Thomas J Gloria, Hermilo R [24 pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 906. |
Effect of Mach number on
boundary-layer transition in the presence of pressure rise and surface
roughness on an ogive-cylinder body with cold wall conditions
(April 20, 1956) by Carros, Robert J [31 pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 907. |
The effect of conical camber on
the static longitudinal, lateral, and directional characteristics of a
45-degree sweptback wing at Mach numbers up to 0.96 (July 03,
1956) by Sammonds, Robert I Reynolds, Robert M [65 pages; 3.6 MB]
|
|
Abstract: No Abstract Available |
| 908. |
An experimental investigation
at Mach numbers from 2.1 to 3.0 of circular-internal-contraction inlets
with translating centerbodies (October 31, 1956) by
Mossman, Emmett A Pfyl, Frank A [29 pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 909. |
An investigation of the lift,
drag, and static-stability characteristics of a triangular-wing
airplane configuration at Mach numbers from 3.00 to 6.28
(December 19, 1956) by Gloria, Hermilo R [18 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 910. |
Aerodynamic characteristics in
pitch of several triple-body missile configurations at Mach numbers 0.6
to 1.4 (November 02, 1956) by Knechtel, Earl D Andrea,
Arvid N [29 pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 911. |
Investigation at supersonic and
subsonic Mach numbers of auxiliary inlets supplying secondary air flow
to ejector exhaust nozzles (January 25, 1956) by Hearth,
Donald P Cubbison, Robert W [48 pages; 1.4 MB] |
|
Abstract: The results indicated increases in
auxiliary-inlet pressure recovery with increases in scoop height
relative to the boundary-layer thickness. The pressure recovery
increased at about the same rate as theoretically predicted for an
inlet in a boundary layer having a one-seventh power profile, but was
only about 0.68 to 0.75 of the theoretically obtainable values. Under
some operating conditions, flow from the primary jet was exhausted
through the auxiliary inlet. This phenomenon could be predicted from
the ejector pumping characteristics. |
| 912. |
Interference effects at Mach
1.9 on a horizontal tail due to trailing shock waves from an
axisymmetric body with an exiting jet (January 25, 1956)
by Salmi, Reino J Klann, John L [36 pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 913. |
Effects of rocket-armament
exhaust gas on the performance of a supersonic-inlet
J34-turbojet-engine installation at Mach 2.0 (February 20,
1956) by Beheim, Milton, A Evans, Phillip J [25 pages; 0.7 MB] |
|
Abstract: No Abstract Available |
| 914. |
External-stream effects on
gross thrust and pumping characteristics of ejectors operating at
off-design Mach numbers (June 26, 1956) by Valerino,
Alfred S Yeager, Richard A [33 pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 915. |
Experimental investigation of
interference effects of lateral-support struts on afterbody pressures
at Mach 1.9 (May 14, 1956) by Klann, John L Huff, Ronald
G [15 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 916. |
Stability of supersonic inlets
at Mach 1.91 with air injection and suction (June 28, 1956)
by Kowalski, K Piercy, Thomas G [36 pages; 1.5 MB] |
|
Abstract: No Abstract Available |
| 917. |
Investigation of the air-flow
regulation characteristics of a translating-spike inlet with two
oblique shocks from Mach 1.6 to 2.0 (July 24, 1956) by
Nettles, J C [16 pages; 0.4 MB] |
|
Abstract: The pressure recovery of an axially
symmetric translating-spike inlet was essentially the same as for a
single cone with the same total angle. In order to match a turbojet
engine over the Mach range of 1.6 to 2.0, the spike translation must be
larger than it is for a single cone. |
| 918. |
Boundary-layer transition at
supersonic speeds (August 03, 1956) by Low, George M [36
pages; 1.1 MB] |
|
Abstract: Recent results of the effects of Mach
number, stream turbulence, leading-edge geometry, leading-edge sweep,
surface temperature, surface finish, pressure gradient, and angle of
attack on boundary-layer transition are summarized. Factors that delay
transition are nose blunting, surface cooling, and favorable pressure
gradient. Leading-edge sweep and excessive surface roughness tend to
promote early transition. The effects of leading-edge blunting on
two-dimensional surfaces and surface cooling can be predicted
adequately by existing theories, at least in the moderate Mach number
range. |
| 919. |
Use of a diffuser Mach number
as a supersonic-inlet control parameter (September 14, 1956)
by Whalen, Paul P Wilcox, Fred A [18 pages; 0.4 MB] |
|
Abstract: A Mach number measured in the subsonic
diffuser was used experimentally as the inlet control parameter of a
bypass control system for an axisymmetric supersonic inlet operated in
combination with a J34 turbo-jet engine at Mach numbers from 1.6 to
2.0. The control maintained the inlet in either critical or
supercritical operation, and, when set for critical diffuser operation,
the control recovered from disturbances that placed the inlet in both
subcritical buzz and supercritical operation. A slotted-rake orifice
gave a more representative value of subsonic diffuser Mach number than
the single total-pressure probe used as a control input. |
| 920. |
Effects of external stream flow
and afterbody variations on the performance of a plug nozzle
(October 02, 1956) by Salmi, R J Cortright, E M , Jr [20 pages;
0.5 MB] |
|
Abstract: The off-design operation of an
isentropic plug nozzle designed for a jet pressure ratio of 15 was
investigated experimentally at subsonic Mach numbers up to 0.9 and jet
pressure ratios up to 5. When installed in a cylindrical nacelle with a
sharp turn at the nozzle lip, the interaction of the jet and the
external stream produced low pressures on the base formed by the high
lip angle. These low pressures increased the nacelle drag and caused an
overexpansion of the jet, which resulted in lower pressures on the plug
and, hence, reduced thrust. With a boattail ahead of the plug nozzle,
the base pressures were increased and the jet overexpansion
significantly reduced. |
| 921. |
Effect of several design
variables on internal performance of convergent-plug exhaust nozzles
(October 29, 1956) by Krull, H George Beale, William T
Schmiedlin, Ralph F [34 pages; 1 MB] |
|
Abstract: Numerous experiments were conducted to
determine factors which affect the internal performance of
convergent-plug exhaust nozzles. The results of these experiments,
which include effects of such things as plug shape, nozzle inlet Mach
number, and outer-shell characteristics, provide a basis for optimum
desifn from the standpoint of weight and size. The results also show
performance penalties which can result when the nozzle is too small. |
| 922. |
Observation of laminar flow on
a blunted 15 degree cone-cylinder in free flight at high reynolds
numbers and free-stream mach numbers to 8.17 (October 15,
1956) by Disher, John H Rabb, Leonard [34 pages; 1 MB] |
|
Abstract: A highly polished 15 degree
included-angle cone-cylinder with hemispherical tip has been flown to
obtain boundary-layer transition and heat-transfer data. The model was
launched from a carrier plane at an altitude of 47,500 feet. Laminar
flow existed at a Reynolds number greater than 38.5 x 10(exp) 6 on the
cylinder when the model was at the peak free-stream Mach number of
8.17. The results indicate an appreciable and favorable effect of tip
bluntness in raising the allowable skin temperature for a given
boundary-layer transition Reynolds number. |
| 923. |
Analysis of limitations imposed
on one-spool turboprop-engine designs by compressors and turbines at
flight mach numbers of 0, 0.6, and 0.8 (December 06, 1956)
by Cavicchi, Richard H [67 pages; 2 MB] |
|
Abstract: Turbine centrifugal stress is a
limiting factor for all flight conditions studied. This stress is more
severe for sea-level operations than for subsonic flight at the
tropopause. Turbines designed for a stress of 30,000 psi are capable of
driving a light, compact, high-spedd compressor but only at high values
of specific fuel consumption. An increase in turbine-inlet temperature
is accompanied by an increase in turbine centrifugal stress. If
stresses in excess of 50,000 psi can be tolerated, compressor
aerodynamics may become a primary limitation. |
| 924. |
Preliminary investigation of
methods to increase base pressure of plug nozzles at Mach 0.9
(December 19, 1956) by Salmi, Reino J [14 pages; 0.4 MB] |
|
Abstract: The effects of various afterbody
changes on the base pressure of a nacelle-type isentropic plug nozzle
installation operating at lower-than-design jet pressure ratios were
investigated at a Mach number of 0.9. Although the estimates of the net
propulsive force contain some uncertainties, the results indicate that
both a plain-ring base shroud and a circular-arc boattail fairing
reduced the loss in net propulsive force experienced with a cylindrical
nacelle installation of the plug nozzle. |
| 925. |
Flight determination of drag
and pressure recovery of two scoop inlets located at
maximum-body-diameter station at Mach numbers from 0.8 to 1.8
(January 10, 1956) by Putland, Leonard W [26 pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 926. |
Pressure distribution and
aerodynamic loadings for several-flap-type trailing-edge controls on a
trapezoidal wing at Mach numbers of 1.61 and 2.01 (March 12,
1956) by Lord, Douglas R Czarnecki, K R [153 pages; 4 MB] |
|
Abstract: No Abstract Available |
| 927. |
Tabulated pressure data for
several flap-type trailing edge controls on a trapezoidal wing at Mach
numbers of 1.61 and 2.01 (February 27, 1956) by Lord,
Douglas R Czarnecki, K R [277 pages; 7.5 MB] |
|
Abstract: No Abstract Available |
| 928. |
A transonic investigation of
changing indentation design Mach number on the aerodynamic
characteristics of a 45 degree sweptback-wing-body combination designed
for high performance (January 10, 1956) by Loving,
Donald L [87 pages; 2.5 MB] |
|
Abstract: No Abstract Available |
| 929. |
Preliminary free-flight study
of the drag and stability of a series of short-span missiles at Mach
numbers from 0.9 to 1.3 (February 08, 1956) by Hall,
James Rudyard [16 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 930. |
The effects at a Mach number of
6.86 of drag brakes on the lift, drag, and pitching moment of an ogive
cylinder (March 19, 1956) by Penland, Jim A Fetterman,
David E , Jr [32 pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 931. |
Aerodynamic loadings associated
with swept and unswept spoilers on a flat-plate at Mach numbers of 1.61
and 2.01 (March 12, 1956) by Lord, Douglas R Czarnecki,
K R [175 pages; 5.3 MB] |
|
Abstract: No Abstract Available |
| 932. |
Investigation of jet effects on
a flat surface downstream of the exit of a simulated turbojet nacelle
at a free-stream Mach number of 1.39 (April 02, 1956) by
Bressette, Walter E Leiss, Abraham [72 pages; 3.4 MB] |
|
Abstract: No Abstract Available |
| 933. |
A free-flight investigation of
the effects of a sonic jet on the total-drag and base-pressure
coefficients of a boattail body of revolution from Mach number 0.83 to
1.70 (March 08, 1956) by Falanga, Ralph A [20 pages; 0.6
MB] |
|
Abstract: No Abstract Available |
| 934. |
Theoretical calculations of the
pressure, forces, and moments at supersonic speeds due to various
lateral motions acting on thin isolated vertical tails (1956)
by Margolis, Kenneth Bobbitt, Percy J [44 pages; 1.6 MB] |
|
Abstract: Velocity potentials, pressure,
distributions, and stability derivatives are derived by use of
supersonic linearized theory for families of thin isolated vertical
tails performing steady rolling, steady yawing, and
constant-lateral-acceleration motions. Vertical-tail families
(half-delta and rectangular plan forms) are considered for a broad Mach
number range. Also considered are the vertical tail with arbitrary
sweepback and taper ratio at Mach numbers for which both the leading
edge and trailing edge of the tail are supersonic and the triangular
vertical tail with a subsonic leading edge and a supersonic trailing
edge. Expressions for potentials, pressures, and stability derivatives
are tabulated. |
| 935. |
Investigation of local
heat-transfer and pressure drag characteristics of a yawed circular
cylinder at supersonic speeds (January 24, 1956) by
Goodwin, Glen Creager, Marcus O Winkler, Ernest L [46 pages; 1.4 MB] |
|
Abstract: Local heat-transfer coefficients,
temperature recovery factors, and pressure distributions were measured
on the front side of a circular cylinder at a nominal Mach number of
3.9 over a range of free-stream Reynolds numbers from 2.1 x 10 to the
3rd power to 6.7 x 10 to the 3rd power and yaw angles from zero degrees
to 44 degrees. Yawing the cylinder reduced the heat-transfer
coefficients and the pressure drag coefficients. The amount of
reduction may be predicted by a theory presented herein. |
| 936. |
Effect of boundary-layer
control and inlet lip shape on the performance of a twin-scoop
air-induction system at Mach numbers from 0 to 1.9 (February
14, 1956) by LAZZERONI FRANK A Pfyl, Frank A [53 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 937. |
Wind-tunnel investigation at
Mach numbers from 0.8 to 1.4 of static longitudinal and
lateral-directional characteristics of an unswept-wing airplane model
(December 13, 1956) by Summers, James L Treon, Stuart L Graham,
Lawrence A [100 pages; 3.4 MB] |
|
Abstract: No Abstract Available |
| 938. |
Free-flight aerodynamic-heating
data to Mach number 10.4 for a modified Von Karman nose shape
(July 10, 1956) by Bland, William M , Jr Collie, Katherine A [29
pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 939. |
Analysis of ram-jet engine
performance including effects of component changes (October
29, 1956) by Weber, Richard J Luidens, Roger W [47 pages; 1.6 MB]
|
|
Abstract: Calculated design-point performance of
ram-jet engines using JP-4 fuel is presented for a wide range of engine
total-temperature ratios and combustion-chamber-inlet Mach numbers for
flight numbers from 1.5 to 4.0. The results include engine thrust,
drag, fuel consumption, and area ratios. Data are also presented to
illustrate the sensitivity of the results to variations in the assumed
component parameters. A brief comparison is included between fixed-and
variable-geometry engines. |
| 940. |
Performance of a blunt-lip side
inlet with ramp bleed, bypass, and a long constant-area duct ahead of
the engine : Mach number 0.66 and 1.5 to 2.1 (December 28,
1956) by Allen, John L [56 pages; 1.6 MB] |
|
Abstract: Unsteady shock-induced separation of
the ramp boundary layer was reduced and stabilized more effectively by
external perforations than by external or internal slots. At Mach 2.0
peak total-pressure recovery was increased from 0.802 to 0.89 and
stable mass-flow range was increased 185 percent over that for the
solid ramp. Peak pressure recovery occurred just before instability.
The 7 and one-third-diameter duct ahead of the engine reduced large
total-pressure distortions but was not as successful for small
distortions as obtained with throat bleed. By removing boundary-layer
air the bypass nearly recovered the total-pressure loss due to the long
duct. |
| 941. |
Investigation of the effects of
body camber and body indentation on the longitudinal characteristics of
a 60 degree delta-wing-body combination at a Mach number of 1.61
(April 20, 1956) by Sevier, John R , Jr [22 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 942. |
Flight investigation of the
effect of a propulsive jet positioned according to the transonic area
rule on the drag coefficients of a single-engine delta-wing
configuration at Mach numbers from 0.83 to 1.36 (April 13,
1956) by Judd, Joseph H Falanga, Ralph A [37 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 943. |
Aerodynamic damping at Mach
numbers of 1.3 and 1.6 of a control surface on a two-dimensional wing
by a free-oscillation method (May 1956) by Tuovila, W J
Hess, Robert W [23 pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 944. |
Measurements of aerodynamic
heat transfer and boundary-layer transition on a 10 degree cone in free
flight at supersonic Mach numbers up to 5.9 (April 26, 1956)
by Rumsey, Charles B Lee, Dorothy B [34 pages; 1.7 MB] |
|
Abstract: No Abstract Available |
| 945. |
Wind-tunnel investigation of a
ram-jet model having a wing and canard surfaces of delta plan form with
70 degrees swept leading edges : force and moment characteristics at
combined angles of pitch and sideslip for Mach number 2.01
(April 26, 1956) by Driver, Cornelius Hamilton, Clyde V [68
pages; 1.6 MB] |
|
Abstract: No Abstract Available |
| 946. |
Aerodynamic characteristics of
a 6-percent-thick symmetrical circular-arc airfoil having a
30-percent-chord trailing-edge flap at a Mach number of 6.9
(May 05, 1956) by Ridyard, Herbert W Fetterman, David E , Jr [50
pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 947. |
Supersonic-area-rule design and
rocket-propelled flight investigation of a zero-lift
straight-wing-body-nacelle configuration between Mach numbers 0.8 and
1.53 (April 26, 1956) by Hoffman, Sherwood [29 pages;
0.9 MB] |
|
Abstract: No Abstract Available |
| 948. |
Turbulent and laminar
heat-transfer measurements on a 1/6-scale NACA RM-10 missile in free
fight to Mach number of 4.2 and to a wall temperature of 1400 R
(July 03, 1956) by Piland, Robert O Collie, Katherine A Stoney,
William E [45 pages; 2.7 MB] |
|
Abstract: No Abstract Available |
| 949. |
Free-flight measurements of the
zero-lift drag of several wings at Mach numbers from 1.4 to 3.8
(June 22, 1956) by Jackson, H Herbert [28 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 950. |
Effect of wing camber and twist
at Mach numbers from 1.4 to 2.1 on the lift, drag, and longitudinal
stability of a rocket-powered model having a 52.5 degree sweptback wing
of aspect ratio 3 and inline tail surfaces (May 07, 1956)
by Gillespie, Warren, Jr [30 pages; 0.7 MB] |
|
Abstract: No Abstract Available |
| 951. |
Jet effects on base and
afterbody pressures of a cylindrical afterbody at transonic speeds
(May 23, 1956) by Cubbage, James M , Jr [51 pages; 1.4 MB] |
|
Abstract: An investigation of the effects of jet
nozzle geometry, size of base annulus, and base bleed upon the base and
afterbody pressures of a cylindrical afterbody at transonic speeds has
been conducted. Sonic and supersonic conical nozzles with jet-to-base
diameter ratios from 0.25 to 0.85 were investigated with a cold jet at
jet total-pressure ratios up to approximately 8.0 through a Mach number
range from 0.6 to 1.25. Base pressure coefficients of about -0.55 were
measured for the sonic nozzles at a Mach number of 1 or greater. The
jet-to-base diameter ratio had a substantial effect on the base
pressure obtained on the cylindrical afterbody of this investigation.
Base bleed was beneficial in increasing the base pressure under certain
conditions but had little or no effect at certain other conditions. |
| 952. |
Heat transfer on the lifting
surfaces of a 60 degree delta wing at angle of attack for Mach number
1.98 (May 31, 1956) by Carter, Howard S [25 pages; 1 MB]
|
|
Abstract: No Abstract Available |
| 953. |
Longitudinal stability
characteristics of a simple infrared homing missile configuration at
Mach numbers of 0.7 to 1.4 (June 12, 1956) by Brown,
Clarence, A , jr [29 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 954. |
Zero-lift drag of a series of
bomb shapes at Mach numbers from 0.60 to 1.10 (July 26, 1956)
by Stoney, William E, Jr Royall, John F [13 pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 955. |
Force and pressure-distribution
measurements at a Mach number of 3.12 of slender bodies having
circular, elliptical, and triangular cross sections and the same
longitudinal distribution of cross-sectional area (July 13,
1956) by Lange, Roy H Wittliff, Charles E [46 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 956. |
Hinge moment and effectiveness
of an unswept constant-chord control and an overhang-balanced, swept
hinge-line control on an 80 degree swept pointed wing at Mach numbers
from 0.75 to 1.96 (August 28, 1956) by Guy, Lawrence D
[40 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
|
Abstract: An investigation has been made to
determine the aerodynamic characteristics of the NACA 4-(5)(05)-041
four-blade, single-relation propeller and the NACA 4-(5)(05)-037 six-
and eight-blade, dual-rotation propellers in combination with various
spinners and NACA d-type spinner-cowling combinations at Mach numbers
up to 0.84. Propeller force characteristics, local velocity
distributions in the propeller planes, inlet pressure recoveries, and
static-pressure distributions on the cowling surfaces were measured for
a wide range of blade angles, advance ratios, and inlet-velocity
ratios. Included are data showing: (a) the effect of extended
cylindrical spinners on the characteristics of the single-rotation
propeller, (b) the effect of variation of the difference in blade angle
setting between the front and rear components of the dual-rotation
propellers, (c) the negative- and static-thrust characteristics of the
propellers with 1 series spinners, and (d) the effects of ideal- and
platform-type propeller-spinner junctures on the pressure-recovery
characteristics of the single-rotation propeller-spinner-cowling
combination. |
| 1010. |
Performance of
external-compression bump inlet at Mach numbers of 1.5 and 2.0
(Apr 1957) by Paul C. Simon, Dennis W. Brown, Ronald G. Huff [40
pages; 1.2 MB] |
|
Abstract: An experimental investigation of a
one-fifth-scale model of the forebody of a proposed supersonic fighter
was conducted to determine the internal performance and configuration
drag of various twin-side inlets. |
| 1011. |
Investigation of a
high-performance top inlet to Mach number of 2.0 and at angles of
attack to 20 degrees (Mar 1957) by Donald J. Vargo,
Philip N. Parks, Owen H. Davis [63 pages; 1.9 MB] |
|
Abstract: Several top-inlet configurations were
tested on a body of revolution in the 8- by 6-foot supersonic wind
tunnel at angles of attack from 0 to 20 degrees and at free-stream Mach
numbers of 1.5 to 2.0. The effect on performance of the following
variable was studied: throat bleed, ramp perforations, inlet approach
surface, side fairing, fuselage fences, canopies, and a simulated 60
degree delta wing. |
| 1012. |
Full-scale free-jet
investigation of a two-shock side-inlet diffuser at Mach 2.75 and a
comparison with a single-shock diffuser (Apr 1957) by
John E. McAulay [24 pages; 0.7 MB] |
|
Abstract: A full-scale free-jet investigation of
a two-shock side-inlet diffuser at a Mach number of 2.75 was conducted
in an NACA Lewis laboratory altitude test chamber. Data were obtained
over ranges of free-stream total pressure and temperature of 3800 to
1930 pounds per square foot and 860 to 990 degrees R, respectively. |
| 1013. |
Aerodynamic performance of
several techniques for spike-position control of a blunt-lip nose inlet
having internal contraction; Mach numbers of 0.63 and 1.5 to 2.0
(Sep 1957) by Arthur A. Anderson, Maynard I. Weinstein [44 pages;
1.2 MB] |
|
Abstract: A study was made to determine locations
of pressure sensors for controlling the spike position of a blunt-lip,
axisymmetric inlet having internal contraction. The inlet performance
was determined at Mach numbers of 0.63 and 1.5 to 2.0 for airflow
schedules corresponding to those of a given turbojet engine over a wide
range of ambient temperatures. |
| 1014. |
Oblique-shock relations at
hypersonic speeds for air in chemical equilibrium (Jan 1957)
by W. E. Moeckel [19 pages; 1 MB] |
|
Abstract: Oblique-shock relations for air in
chemical equilibrium have been calculated for flight velocities up to
25,000 feet per second at altitudes up to 200,000 feet. Results show
that those shock parameters which are functions only of Mach number
normal to the shock for an ideal gas are strongly influenced by flight
altitude (initial conditions), as well as normal Mach number, when
dissociation takes place. |
| 1015. |
Heat transfer and
boundary-layer transition on two blunt bodies at Mach numbers 3.12
(Oct 1957) by N. S. Diaconis, Richard J. Wisniewski, John R. Jack
[32 pages; 0.8 MB] |
|
Abstract: Local heat-transfer parameters were
measured on a hemisphere-cone-cylinder and on a 120
degree-included-angle cone-cylinder at a free-stream Mach number of
3.12 and at free-stream static temperature. |
| 1016. |
On the minimization of airplane
responses to random gusts (Oct 1957) by Murray Tobak [72
pages; 2.1 MB] |
|
Abstract: A theoretical study is made of the
motions experienced by aircraft in response to sharp-edge, harmonic,
and random gusts. For the sharp-edge and harmonic gusts, exact
responses in normal acceleration and pitching velocity are presented
for the rectangular wing flying at Mach number 1.2. |
| 1017. |
Effect of ambient-temperature
variation on the matching requirements of inlet-engine combinations at
supersonic speeds (Jan 1957) by Eugene Perchonok, Donald
P. Hearth [17 pages; 0.6 MB] |
|
Abstract: The effect of ambient temperature on
the matching requirements of inlet-engine combinations has been
analyzed for two typical turbojet engines up to a Mach number of 3.5.
The changes in ambient temperature ordinarily encountered in flight can
markedly influence the performance of matched inlet-engine combinations
for engines operated at constant mechanical speed. |
| 1018. |
Comparison of experimental and
theoretical zero-lift wave-drag results for various wing-body-tail
combinations at Mach numbers up to 1.9 (March 27, 1957)
by Peterson, Robert B [49 pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 1019. |
Reductions in
temperature-recovery factor associated with pulsating flows generated
by spike-nosed cylinders at a Mach number of 3.50 (March 04,
1957) by Hermach, C A Kraus, Samuel Reller, John O , Jr [26
pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 1020. |
Lateral-directional aerodynamic
characteristics of several coplanar triple-body missile configurations
at Mach numbers from 0.6 to 1.4 (April 10, 1957) by
Treon, Stuart L Knechtel, Earl D [28 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 1021. |
Effects of string-support
interference on the drag of an olgive-cylinder body with and without a
boatail at 0.6 to 1.4 Mach number (December 03, 1957) by
Lee, George Summers, James L [29 pages; 0.7 MB] |
|
Abstract: No Abstract Available |
| 1022. |
The effects of boundary-layer
separation over bodies of revolution with conical tail flares
(December 12, 1957) by Dennis, David H [36 pages; 1.1 MB] |
|
Abstract: The magnitude and the effects of
boundary-layer separation on normal-force-curve slopes, centers of
pressure, pressure distributions, and lift and drag coefficients were
determined for various bodies of revolution with conical tail flares at
Mach numbers from 3.0 to 6.3. Some of the experimental results are
compared to theoretical predictions of the aerodynamic characteristics
of the bodies. |
| 1023. |
Investigation of combustion in
16-inch ram jet under simulated conditions of high altitude and high
Mach number (June 27, 1957) by Nussdorfer, T J
Sederstrom, D C Perchonok, E [54 pages; 1.4 MB] |
|
Abstract: Results obtained with three different
burner configurations in a connected-pipe investigation of a 16-inch
ram jet are presented. The radial position of the fuel injector and the
engine-outlet area both affected burner performance. For a given
configuration, only slight changes in total-pressure ratio across the
combustion chamber were obtained over the complete range of operation.
With one burner, combustion efficiencies obtained at a
combustion-chamber-inlet static pressure of 800 pounds per square foot
absolute were greater than those obtained at 1920 pounds per square
foot absolute. |
| 1024. |
Experimental investigation of
water injection in subsonic diffuser of a conical inlet operation at
free-stream Mach number of 2.5 (January 15, 1957) by
Beke, Andrew [12 pages; 0.4 MB] |
|
Abstract: A spike-type nose inlet with sharp-lip
cowl was investigated at a free-stream Mach number of 2.5 with water
injection in its 16-inch diameter, 11-foot-long subsonic diffuser
section. Inlet total temperature of exit with liquid-air ratios of
about 0.04 with no apparent change in the critical pressure recovery.
The observed temperature drops were less than the theoretically
predicted values, and the amount of water evaporated was 35 to 50
percent less than that theoretically possible. |
| 1025. |
Internal performance of several
auxiliary air inlets immersed in a turbulent boundary layer at Mach
numbers of 1.3, 1.5, and 2.0 (January 18, 1957) by Huff,
Ronald G Anderson, Arthur R [25 pages; 0.7 MB] |
|
Abstract: Internal performance of normal-shock
rectangular, circular, and scoop inlets and of external-compression
inlets experimentally obtained with varying immersion in a turbulent
boundary layer. Recoveries varied from about 95 percent of theoretical
in the free stream to 80 percent with complete immersion, while the
corresponding mass flows were usually above 95 percent of theoretical.
Turning of the flow through 10 degrees caused losses in pressure
recovery of 0.03 to 0.07. External compression did not improve pressure
recovery in the boundary layer. Average distortion at critical
operation for all inlets was 5 percent. |
| 1026. |
Comparison of effect of a
turbojet engine and three cold-flow configurations on the stability of
a full-scale supersonicle inlet (January 24, 1957) by
Musial, Norman T [17 pages; 0.5 MB] |
|
Abstract: Increasing the volume and length of the
duct behind the inlet affected the inlet stability at Mach 2.0 and zero
angle of attack. Close approximation of the inlet stability limit of
the J34 engine-inlet configuration was obtained by a cold-pipe
configuration having a length and volume approaching that measured to
the engine turbine. Variation of these parameters had a small effect on
the minimum subcritical stable mass flow below a cowl-lip-position
parameter of 44 degrees and appeared to have a negligible effect on the
inlet pressure-recovery - mass-flow curve. Initial buzz frequency and
minimum cowl-lip-position parameter for complete buzz-free operation
varied with configuration. |
| 1027. |
An inlet design concept to
reduce flow distortion at angle of attack (February 26, 1957)
by Schueller, Carl F Stitt, Leonard E [24 pages; 1.1 MB] |
|
Abstract: Flow distortions were measured at the
inlet face and diffuser exit of three axisymmetric inlets operating at
angles of attack of 0 degree to 14 degrees and at a Mach number of
1.91. |
| 1028. |
Investigation of mass-flow and
pressure recovery characteristics of several underslung scoop-type
inlets at free-stream Mach numbers of 2.0, 1.8, 1.5, and 0.66
(March 13, 1957) by Valerino, Alfred S Zappa, Robert F [40 pages;
1 MB] |
|
Abstract: No Abstract Available |
| 1029. |
Observation of laminar flow on
an air-launched 15 degree cone-cylinder at local Reynolds numbers to 50
x 10(exp 6) at peak Mach number of 6.75 (March 04, 1957)
by Rabb, Leonard Krasnican, Milan J [34 pages; 1 MB] |
|
Abstract: No Abstract Available |
| 1030. |
Performance of a supersonic
ramp-type side inlet with ram-scoop throat bleed and varying fuselage
boundary-layer removal : Mach number range 1.5 to 2.0 / Glenn A.
Mitchell and Robert C. Campbell (January 17, 1957) by
Mitchell, Glenn A Campbell, Robert C [30 pages; 1 MB] |
|
Abstract: Provided sufficient throat bleed was
employed, maximum pressure recoveries of 0.87 to 0.88 at Mach number
2.0 were obtained for a fuselage-mounted 14 degrees ramp inlet
regardless of the amount of fuselage boundary layer ingested. The
addition of inlet side fairings yielded further increases in pressure
recovery to 0.90 to 0.91, decreased critical drag coefficients, and
increased critical mass-flow ratios. With throat bleed, peak pressure
recoveries and calculated thrust-minus-drag values were comparable at
two axial positions of the scoop and were highest with the greatest
amount of fuselage boundary layer ingested. |
| 1031. |
Investigation of
shock-boundary-layer interaction on the spike of a conical-spike nose
inlet (January 09, 1957) by Wise, George A Sterbentz,
William H [19 pages; 0.8 MB] |
|
Abstract: Measurements were made of the height of
the shock-induced boundary-layer thickening and separation over a Mach
number range of 1.6 to 2.0. The behavior of the interaction depended on
longitudinal spike position as well as on cone surface Mach number. The
cone position affected the interaction by changing the rate of subsonic
diffusion and thereby changing the pressure aft of the terminal shock.
When the pressure rise due to the interaction exceeded about 1.9, the
boundary layer was separated. |
| 1032. |
Some operating experience and
problems encountered during operation of a free-jet facility
(February 13, 1957) by Mcaulay, John E Prince, William R [22
pages; 0.7 MB] |
|
Abstract: During a free-jet investigation of a
28-inch ram-jet engine at a Mach number of 2.35, flow pulsation at the
engine inlet were discovered which proved to have an effect on the
engine performance and operational characteristics, particularly the
engine rich blowout limits. This report discusses the finding of the
flow pulsations, their elimination, and effect. Other facility
characteristics, such as the establishment of flow simulation and the
degree of subcritical operation of the diffuser, are also explained. |
| 1033. |
Investigation of a
supersonic-inlet - turbojet-engine combination at Mach 2.0 and angles
of attack up to 6 degrees (July 1957) by Hearth, Donald
P Musial, Norman T [26 pages; 0.7 MB] |
|
Abstract: No Abstract Available |
| 1034. |
Jet effects on base pressures
of conical afterbodies at Mach 1.91 and 3.12 (August 12, 1957)
by Baughman, L Eugene Kochendorfer, Fred D [113 pages; 5.6
MB] |
|
Abstract: No Abstract Available |
| 1035. |
Exploratory investigation of
aerodynamic effects of external combustion of aluminum borohydride in
airstream adjacent to flat plate in Mach 2.46 tunnel (July 29,
1957) by Dorsch, Robert G Serafini, John S Fletcher, Edward A [92
pages; 3.4 MB] |
|
Abstract: No Abstract Available |
| 1036. |
Free-flight determination of
boundary-layer transition and heat transfer for a hemisphere-cylinder
at Mach numbers to 5.6 (October 21, 1957) by Krasnican,
M J Wisniewski, R J [46 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 1037. |
Total-pressure distortion and
recovery of supersonic nose inlet with conical centerbody in subsonic
icing conditions (September 17, 1957) by Gelder, Thomas
F [42 pages; 1.5 MB] |
|
Abstract: Ice was formed on a full-scale unheated
supersonic nose inlet in the NACA Lewis icing tunnel to determine its
effect on compressor-face total-pressure distortion and recovery.Inlet
angle of attack was varied from 0degrees to 12 degrees, free-stream
Mach number from 0.17 to 0.28, and compressor-face Mach number from
0.10 to 0.47. Icing-cloud liquid-water content was varied from 0.65 to
1.8 grams per cubic meter at free-stream static air temperatures of 15
degrees and 0 degrees F. The addition of ice to the inlet components
increased total-pressure-distortion levels and decreased recovery
values compared withclear0air results, the losses increasing with time
in ice. The combination of glaze ice, high corrected weight flow, and
high angle of attack yielded the highest levels of distortion and
lowest values of recovery. The general character of compressor-face
distortion with an iced inlet was the same as that for the clean inlet,
the total-pressure gradients being predominantly radial, with
circumferential gradients occurring at angle of attack. |
| 1038. |
Pressure drag of axisymmetric
cowls having large initial lip angles at Mach numbers from 1.90 to 3.88
(October 21, 1957) by Samanich, Nick E [17 pages;
0.9 MB] |
|
Abstract: No Abstract Available |
| 1039. |
Performance of a
translating-double-cone axisymmetric inlet with cowl bypass at Mach
numbers from 2.0 to 3.5 (November 13, 1957) by Connors,
James F Wise, George A [26 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 1040. |
A reexamination of the use of
simple concepts for predicting the shape and location of detached shock
waves (December 1957) by Love, Eugene S [54 pages; 1.7
MB] |
|
Abstract: A reexamination has been made of the
use of simple concepts for predicting the shape and location of
detached shock waves. The results show that simple concepts and
modifications of existing methods can yield good predictions for many
nose shapes and for a wide range of Mach numbers. |
| 1041. |
Some experimental studies of
panel flutter at Mach number 1.3 (February 1957) by
Sylvester, Maurice A Baker, John E [26 pages; 0.8 MB] |
|
Abstract: Experimental studies of panel flutter
using thin metal plates were conducted at a Mach number of 1.3 to
verify its existence and to study the effects of some structural
parameters on the flutter characteristics. The effects of tensile
forces and buckling were studied on panels clamped front and rear, in
addition to initially buckled panels clamped on all four edges. Panel
flutter was obtained under controlled laboratory conditions and it was
found that tensile forces, shortening the panels, and increasing the
bending stiffness were effective means for eliminating flutter. Buckled
panels were more susceptible to flutter than unbuckled panels. No
apparent systematic trends in the flutter modes or frequencies could be
observed. |
| 1042. |
Base pressure at supersonic
speeds on two-dimensional airfoils and on bodies of revolution with and
without fins having turbulent boundary layers (1957) by
LOVE EUGENE S [66 pages; 2.3 MB] |
|
Abstract: An analysis has been made of available
experimental data to show the effects of most of the variables that are
more predominant in determining base pressure at supersonic speeds. The
analysis covers base pressures for two-dimensional airfoils and for
bodies of revolution with and without stabilizing fins and is
restricted to turbulent boundary layers. The present status of
available experimental information is summarized as are the existing
methods for predicting base pressure. A simple semiempirical method is
presented for estimating base pressure. For two-dimensional bases, this
method stems from an analogy established between the base-pressure
phenomena and the peak pressure rise associated with the separation of
the boundary layer. An analysis made for axially symmetric flow
indicates that the base pressure for bodies of revolution is subject to
the same analogy. Based upon the methods presented, estimations are
made of such effects as Mach number, angle of attack, boattailing,
fineness ratio, and fins. These estimations give fair predictions of
experimental results. (author) |
| 1043. |
Experimental investigation of
the forces and moments due to sideslip of a series of triangular
vertical- and horizontal-tail combinations at Mach numbers of 1.62,
1.93, and 2.41 (March 1957) by Coletti, Donald E [33
pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 1044. |
Investigation of downwash,
sidewash, and Mach number distribution behind a rectangular wing at a
Mach number of 2.41 (1957) by Adamson, David Boatright,
William B [58 pages; 3 MB] |
|
Abstract: An investigation of the nature of the
flow field behind a rectangular wing of circular arc cross section has
been conducted in the Langley 9-inch supersonic tunnel. Pitot- and
static-pressure surveys covering a region of flow behind the wing have
been made together with detailed pitot surveys throughout the region of
the wake. In addition, the flow direction has been measured by means of
a weathercocking vane. Theoretical calculations have been made to
obtain the variation of both downwash and sidewash with angle of attack
by using the superposition method of Lagerstrom, Graham, and
Grosslight. In addition, the effect of wing thickness on the sidewash
with the wing at 0 degree angle of attack has been evaluated. |
| 1045. |
Temperature measurements from a
flight test of two wing-body combinations at 7 degree angle of attack
for Mach numbers to 4.86 and Reynolds numbers to 19.2 X 10(exp 6)
(September 12, 1957) by Chauvin, Leo T [37 pages; 2.7 MB] |
|
Abstract: No Abstract Available |
| 1046. |
Pressure distribution induced
on a flat plate by a supersonic and sonic jet exhaust at a free-stream
Mach number of 1.80 (January 10, 1957) by Leiss, Abraham
Bressette, Walter E [62 pages; 2.7 MB] |
|
Abstract: No Abstract Available |
| 1047. |
Flight investigation of a ram
jet burning magnesium slurry fuel and having a conical shock inlet
designed for a Mach number of 4.1 (January 22, 1957) by
Barlett, Walter A , Jr Merlet, Charles F [24 pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 1048. |
Aerodynamic forces and moments
on a large ogive-cylinder store at various locations below the fuselage
center line of a swept-wing bomber configuration at a Mach number of
1.61 (January 14, 1957) by Morris, Odell A [45 pages;
1.1 MB] |
|
Abstract: No Abstract Available |
| 1049. |
Experimental static aerodynamic
forces and moments at high subsonic speeds on a missile model during
simulated launching from the midsemispan location of a 45 degree
sweptback wing-fuselage-pylon combination (January 10, 1957)
by Alford, William J King, Thomas, Jr [48 pages; 2.9 MB] |
|
Abstract: An investigation was made at high
subsonic speeds in the Langley high-speed 7- by 10-foot tunnel to
determine the static aerodynamic forces and moments on a missile model
during simulated launching from the midsemispan location of a 45 degree
sweptback wing-fuselage-pylon combination. The results indicated
significant variations in all the aerodynamic components with changes
in chordwise location of the missile. Increasing the angle of attack
caused increases in the induced effects on the missile model because of
the wing-fuselage-pylon combination. Increasing the Mach number had
little effect on the variations of the missile aerodynamic
characteristics with angle of attack except that nonlinearities were
incurred at smaller angles of attack for the higher Mach numbers. The
effects of finite wing thickness on the missile characteristics, at
zero angle of attack, increase with increasing Mach number. The effects
of the pylon on the missile characteristics were to causeincreases in
the rolling-moment variation with angle of attack and a negative
displacement of the pitching-moment curves at zero angle of attack. The
effects of skewing the missile in the lateral direction relative to and
sideslipping the missile with the wing-fuselage-pylon combination were
to cause additional increments in side force at zero angle of attack.
For the missile yawing moments the effects of changes in skew or
sideslip angles were qualitatively as would be expected from
consideration of the isolated missile characteristics, although there
existed differences in theyawing-moment magnitudes. |
| 1050. |
Hinge-moment and effectiveness
characteristics of an aspect-ratio-8.2 flap-type control on a 60 degree
delta wing at Mach numbers from 0.72 to 1.96 (January 07,
1957) by Guy, Lawrence D [55 pages; 1.6 MB] |
|
Abstract: No Abstract Available |
| 1051. |
Drag of conical and
circular-arc boattail afterbodies at Mach numbers from 0.6 to 1.3
(January 22, 1957) by Silhan, Frank V Cubbage, James M , Jr [41
pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 1052. |
Zero-lift drag of a large
fuselage cavity and a partially submerged store on a 52.5 degree
sweptback-wing-body configuration as determined from free-flight tests
at Mach numbers of 0.7 to 1.53 (February 26, 1957) by
Hoffman, Sherwood [25 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 1053. |
Stability of two
rocket-propelled models having aspect-ratio-5 unswept tails on a long
body for the Mach number range of 1.7 to 2.4 (March 27, 1957)
by Lundstrom, Reginald R [40 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 1054. |
Hinge-moment characteristics
for a series of controls and balancing devices on a 60 degree delta
wing at Mach numbers of 1.61 and 2.01 (April 12, 1957)
by Lord, Douglas R Czarnecki, K R [69 pages; 1.7 MB] |
|
Abstract: No Abstract Available |
| 1055. |
Measurement of aerodynamic heat
transfer to a deflected trailing-edge flap on a delta fin in free
flight at Mach numbers from 1.5 to 2.6 (April 10, 1957)
by Chauvin, Leo T Buglia, James J [19 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 1056. |
Jet effects on the drag of
conical afterbodies for Mach numbers of 0.6 to 1.28 (April 12,
1957) by Cubbage, James M , Jr [64 pages; 2.1 MB] |
|
Abstract: No Abstract Available |
| 1057. |
A flight investigation to
determine the effectiveness of Mach number 1.0, 1.2, and 1.41 fuselage
indentations for reducing the pressure drag of a 45 degree sweptback
wing configuration at transonic and low supersonic speeds (May
16, 1957) by Blanchard, Willard S , Jr Hoffman, Sherwood [25
pages; 0.9 MB] |
|
Abstract: No Abstract Available |
| 1058. |
Some effects of heat transfer
at Mach number 2.0 at stagnation temperatures between 2,310 and 3,500 R
on a magnesium fin with several leading-edge modifications
(April 18, 1957) by Bland, William M , Jr Bressette, Walter E [30
pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 1059. |
Heat-transfer and pressure
distribution on six blunt noses at a Mach number of 2 (April
18, 1957) by Carter, Howard S Bressette, Walter E [27 pages; 1.4
MB] |
|
Abstract: No Abstract Available |
| 1060. |
Heat transfer and
boundary-layer transition on a highly polished hemisphere-cone in free
flight at Mach numbers up to 3.14 and Reynolds numbers up to 24 x
10(exp 6) (April 18, 1957) by Buglia, James J [27 pages;
1.1 MB] |
|
Abstract: No Abstract Available |
| 1061. |
Static longitudinal and lateral
stability parameters of three flared-skirt two-stage missile
configurations at a Mach number of 6.86 (June 05, 1957)
by Penland, Jim A Carroll, C Maria [49 pages; 4.9 MB] |
|
Abstract: No Abstract Available |
| 1062. |
Preliminary results from a
free-flight investigation of boundary-layer transition and heat
transfer on a highly polished 8-inch-diameter hemisphere-cylinder at
Mach numbers up to 3 and Reynolds numbers based on a length of 1 foot
up to 17.7 x 10(exp (May 16, 1957) by Hall, James R
Speegle, Katherine C Piland, Robert O [28 pages; 1.7 MB] |
|
Abstract: No Abstract Available |
| 1063. |
Aerodynamic characteristics of
missile configurations with wings of low aspect ratio for various
combinations of forebodies, afterbodies, and nose shapes for combined
angles of attack and sideslip at a Mach number of 2.01 (June
25, 1957) by Robinson, Ross B [215 pages; 24.8 MB] |
|
Abstract: An investigation has been made in the
Langley 4-by-4-foot supersonic pressure tunnel to determine the
aerodynamic characteristics of a series of missile configurations
having low-aspect-ratio wings at a Mach number of 2.01. The effects of
wing plan form and size, length-diameter ratio, forebody and afterbody
length, boattailed and flared afterbodies, and component force and
moment data are presented for combined angles of attack and sideslip to
about 28 degrees. No analysis of the data was made in this report. |
| 1064. |
Effects of wing inboard
plan-form modifications on lift, drag, and longitudinal stability at
Mach numbers from 1.0 to 2.3 of a rocket-propelled free-flight model
with a 52.5 degree sweptback wing of aspect ratio 3 (June 19,
1957) by Henning, Allen B [25 pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 1065. |
Investigation at Mach numbers
from 0.80 to 1.43 of pressure and load distributions over a thin 45
degree sweptback highly tapered wing in combination with basic and
indented bodies (June 28, 1957) by Fischetti, Thomas L
[95 pages; 3.6 MB] |
|
Abstract: No Abstract Available |
| 1066. |
Experimental and theoretical
aerodynamic characteristics of two low-aspect-ratio delta wings at
angles of attack to 50 degrees at a Mach number of 4.07 (July
10, 1957) by Smith, Fred M [28 pages; 0.8 MB] |
|
Abstract: No Abstract Available |
| 1067. |
Limited heat-transfer, drag,
and stability results from an investigation at Mach numbers up to 9 of
a large rocket-propelled 10 degree cone (July 22, 1957)
by Hall, James R Speegle, Katherine C [27 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 1068. |
The aerodynamic characteristics
of a body in the two-dimensional flow field of a circular-arc wing at a
Mach number of 2.01 (July 02, 1957) by Gapcynski, John P
Carlson, Harry W [50 pages; 3.2 MB] |
|
Abstract: No Abstract Available |
| 1069. |
Aerodynamic heating and
boundary-layer transition on a 1/10-power nose shape in free flight at
Mach numbers up to 6.7 and free-stream Reynolds numbers up tp 16 x
10(exp 6) (June 17, 1957) by Garland, Benjamine J
Swanson, Andrew G Speegle, Katherine C [32 pages; 1.6 MB] |
|
Abstract: No Abstract Available |
| 1070. |
Rocket-model investigation of
hinge-moments on a trailing-edge control on a 52.5 degree swept wing
between Mach numbers of 0.70 and 1.80 (August 12, 1957)
by Martz, C William [36 pages; 1.1 MB] |
|
Abstract: No Abstract Available |
| 1071. |
Two-dimensional airfoil
characteristics of four NACA 6A-series airfoils at transonic Mach
numbers up to 1.25 (August 06, 1957) by Ladson, Charles
L [47 pages; 1.3 MB] |
|
Abstract: A two-dimensional wind-tunnel
investigation of the flow and force characteristics of four NACA
6A-series airfoils with thickness ratios of 4, 6, and 9 percent has
been conducted in the Langley airfoil test apparatus at at transonic
Mach numbers between 0.8 and 1.25. The Reynolds number range for these
tests varied from 2.6 x 10(6) to 2.8 x 10(6). As was expected, the
airfoils exhibited a smooth transition in force coefficients from a
Mach number of 1.0 to the values obtained at the higher speeds.
Lift-curve slope and maximum lift-drag ratio correlated very well on a
basis of the transonic similarity laws at Mach numbers above 1.0, but
below that value the correlation was not good. The measured effect of
thickness on the drag coefficient at supersonic speeds was less than
that predicted by the transonic similarity laws. Good correlation of
the drag coefficients was obtained by reducing the exponent of the
thickness term from the theoretical value of 1.67 to 1.50. This change
did not affect the correlation at subsonic speeds, which was good for
either case. |
| 1072. |
Aerodynamic heating of a thin,
unswept, untapered, multiweb, aluminum-alloy wing at Mach numbers up to
2.67 as determined from a free-flight investigation of a
rocket-propelled model (August 06, 1957) by Strass, H
Kurt Stephens, Emily W [55 pages; 3.1 MB] |
|
Abstract: No Abstract Available |
| 1073. |
Free-flight aerodynamic-heating
data to a Mach number of 15.5 on a blunted conical nose with a total
angle of 29 degrees (August 1957) by Bland, William M ,
Jr Rumsey, Charles B Lee, Dorothy B Kolenkiewicz, Ronald [43 pages; 2.1
MB] |
|
Abstract: No Abstract Available |
| 1074. |
Comparison of low-lift drag at
Mach numbers from 0.74 to 1.37 of rocket-boosted models having
externally braced wings and cantilever wings (September 25,
1957) by Dickens, Waldo L Hastings, Earl C , Jr [23 pages; 0.7
MB] |
|
Abstract: No Abstract Available |
| 1075. |
Experimental determination of
damping in pitch of swept and delta wings at supersonic Mach numbers
(September 19, 1957) by Moore, John A [24 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 1076. |
Effects of wing warp on the
lift, drag, and static longitudinal stability characteristics of an
aircraft configuration having an arrow wing of aspect ratio 1.86 at
Mach numbers from 1.1 to 1.7 (August 30, 1957) by
Gillespie, Warren, Jr [28 pages; 0.7 MB] |
|
Abstract: No Abstract Available |
| 1077. |
Effect of conical and flat
sting-mounted windshields on the zero-lift drag of a flare-stabilized
bluff body at Mach numbers from 0.6 to 1.15 (September 12,
1957) by Blanchard, Willard S [9 pages; 0.2 MB] |
|
Abstract: No Abstract Available |
| 1078. |
Free-flight investigation of
the drag of a model of a 60 degree delta-wing bomber with strut-mounted
siamese nacelles and indented fuselage at Mach numbers from 0.80 to
1.35 (September 25, 1957) by Hoffman, Sherwood [41
pages; 1.3 MB] |
|
Abstract: No Abstract Available |
| 1079. |
Free-flight skin-temperature
and surface-pressure measurements on a highly polished nose having a
100 degree total-angle cone and a 10 degree half-angle conical flare
section up to a Mach number of 4.08 (August 23, 1957) by
Rashis, Bernard Bond, Aleck C [24 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 1080. |
Tests of aerodynamically heated
multiweb wing structures in a free jet at Mach number 2 : two
aluminum-alloy models of 20-inch chord with 0.064-inch-thick skin at
angles of attack of 0 degree and plus or minus 2 degrees
(October 28, 1957) by Miltonberger, Georgene H Davidson, John R
Griffith, George E [37 pages; 1.3 MB] |
|
Abstract: No Abstract Available |
| 1081. |
Tabulated pressure data for a
series of controls on a 40 degree sweptback wing at Mach numbers of
1.61 and 2.01 (November 08, 1957) by Lord, Douglas R
[336 pages; 13.1 MB] |
|
Abstract: No Abstract Available |
| 1082. |
Aerodynamic load distribution
over a 45 degree swept wing having a spoiler-slot-deflector aileron and
other spoiler ailerons for Mach numbers from 0.60 to 1.03
(December 05, 1957) by West, F E , Jr Whitcomb, Charles F
Schmeer, James W [119 pages; 6.7 MB] |
|
Abstract: No Abstract Available |
| 1083. |
Effects of inlet modification
and rocket-rack extension on the longitudinal trim and low-lift drag of
the Douglas F5D-1 airplane as abtained with a 0.125-scale
rocket-boosted model between mach numbers of 0.81 and 1.64
(1957) by Hastings, Earl C. Jr., Dickens, Waldo L. [27 pages; 1.2
MB] |
|
Abstract: (abstract not available) |
| 1084. |
The effect of body contouring
on the longitudinal characteristics at Mach numbers up to 0.92 of a
wing-fuselage-tail and several wing-fuselage combinations having
sweptback wings of relatively high aspect ratio (1957)
by Sutton, Fred B., Lautenberger, J. Walter, Jr. [31 pages; 2 MB] |
|
Abstract: An investigation was conducted in the
Ames 12 foot pressure wind tunnel to determine the effect of a
realtively simple Kuchemann type fuselage modification at the
wing-fuselage juncture of several sweptback wing-fuselage and
wing-fuselage-tail combinations. |
| 1085. |
Effects of wing-tip droop on
the longitudinal characteristics of two highly swept wing-body
combinations at Mach numbers from 0.6 to 1.4 (1957) by
Knechtel, Earl D., Lee, George. [27 pages; 1.1 MB] |
|
Abstract: Longitudinal aerodynamic
characteristics at Mach numbers from 0.6 to 1.4 were measured to show
the effect of drooping the outboard portions of wings having 53 degrees
and 63 degrees of leading-edge sweep. |
| 1086. |
Effects of vertical location of
wing and horizontal tail on the aerodynamic characteristics in pitch at
Mach numbers from 0.60 to 1.40 of an airplane configuration with an
unswept wing (1957) by Stivers, Louis S., Jr., Lippmann,
Garth W. [58 pages; 2.3 MB] |
|
Abstract: An investigation was conducted in the
Ames 2- by2-foot transonic wind tunnel of a wing-body-tail combination.
|
|
Abstract: A reevaluation of existing flight data
obtained for a supersonic canard missile configuration of the
boost-glide type has been made to determine derivatives not previously
evaluated. These derivatives, together with those previously evaluated
and published, have been utilized to determine some typical airframe
frequency responses based on the three-degree-of-freedom longitudinal
equations of motion. For constant flight conditions, it is shown that
besides the usual drag-coefficient variation, the velocity derivatives
also vary with angle of attack or Mach number. The effect of
angle-of-attack variation on the missile frequency responses which
include the velocity derivatives in their solution is appreciable only
in the low-frequency region of operation (below 10 radians/sec). When
the velocity derivatives are neglected, this low-frequency variation
with angle of attack is different. Some of these differences are
pointed out in the results to indicate the significance of including
the velocity derivatives in the solution for the transfer functions. |
| 1121. |
Shape of initial portion of
boundary of supersonic axisymmetric free jets at large jet pressure
ratios (Jan 1958) by Eugene S. Love, Louise P. Lee [30
pages; 1.1 MB] |
|
Abstract: Calculations have been made of the
initial portion of the boundary of axisymmetric free jets exhausting at
large pressure ratios from a conically divergent nozzle having a jet
exit Mach number of 2.5 and a semidivergence angle of 15 degrees. The
results of the calculations indicate the size and shape of the jet to
be expected at large pressure ratios, the effects of ratio of specific
heats, and the large initial inclinations of the boundary that are
likely to be encountered by hypersonic vehicles at high altitude. |
| 1122. |
Collection of zero-lift drag
data on bodies of revolution from free-flight investigations
(Jan 1958) by William E. Stoney, Jr. [374 pages; 32.4 MB] |
|
Abstract: This report presents a compilation of
most of the zero-lift drag results obtained from free-flight
measurements made by the Langley Pilotless Aircraft Research Division
on fin-stabilized bodies of revolution. The data are arranged on
standard forms, which also contain the significant geometrical factors.
Supplementary data have been provided to facilitate the determination
of the body pressure drags from the measured total drags. Summary plots
and discussions have been included to provide a unified and broad
picture of the effects of body geometry on zero-lift drag. The Mach
number range of the tests extends from 0.6 to approximately 2.0 and the
Reynolds numbers based on body length from 2 x 10-to-the-sixth to 100 x
10-to-the-sixth. |
| 1123. |
Compilation of information on
the transonic attachment of flows at the leading edges of airfoils
(Feb 1958) by Walter F. Lindsey, Emma Jean Landrum [64 pages; 2.4
MB] |
|
Abstract: Schlieren photographs have been
compiled of the two-dimensional flow at transonic speeds past 37
airfoils. These airfoils have variously shaped profiles, and some are
related in thickness and camber. The data for these airfoils were
analyzed to provide basic information on the flow changes involved and
to determine factors affecting transonic-flow attachment, which is a
transition from separated to unseparated flow at the leading edges of
two-dimensional airfoils at fixed angles as the subsonic Mach number is
increased. |
| 1124. |
Turbulent boundary layer on a
yawed cone in a supersonic stream (Jan 1958) by Willis
H. Braun [40 pages; 1.2 MB] |
|
Abstract: The momentum integral equations are
derived for the boundary layer on an arbitrary curved surface, using a
streamline coordinate system. Computations of the turbulent boundary
layer on a slightly yawed cone are made for a Prandtl number 0.70, wall
to free-stream temperature ratios of 1/2, 1, and 2, and Mach numbers
from 1 to 4. Deflection of the fluid in the boundary layer from outer
stream direction, local friction coefficient, displacement surface,
lift coefficient, and pitching-moment coefficient are presented. |
| 1125. |
Recovery temperatures and heat
transfer near two-dimensional roughness elements at Mach 3.1
(Feb 1958) by Paul F. Brinich [21 pages; 0.8 MB] |
|
Abstract: An investigation was made to determine
the effect of single and multiple two-dimensional roughness elements on
the temperature distribution, the pressure distribution, and the heat
transfer at Mach 3.1. A hollow cylinder and a cone-cylinder model were
used. Abrupt perturbations in surface temperature occurred in the
neighborhood of the elements when the boundary layer was turbulent, but
were absent when it was laminar. The type of perturbation depended on
the element shape, forward-facing wedges giving the lowest temperatures
immediately behind the element and forward-facing steps the highest.
For a turbulent boundary layer the heat-transfer rate behind the wedge
element was less than that obtained immediately ahead of the element. |
| 1126. |
Tables and graphs of
normal-shock parameters at hypersonic Mach numbers and selected
altitudes (September 1958) by Paul W. Huber [27 pages;
3.9 MB] |
|
Abstract: Tables and graphs of normal-shock
parameters are presented for real air in thermal and chemical
equilibrium at conditions ahead of the shock corresponding to six
selected altitudes, and for temperatures behind the shock from 2,000
deg K to 11,000 deg K. The altitudes used are those representing the
boundaries of the isothermal layers in that part of the earth's
atmosphere considered applicable to aerodynamic flight; that is, below
an altitude of 300,000 feet. The altitude data and the real-air
thermodynamic data used are reliable for application to this range of
altitudes. Tabulated values at each altitude as a function of the
temperature behind the the shock are presented to show the variation of
the normal-shock parameters with flight Mach number and altitude, and
some discussion of the dependence of the parameters on the initial
pressure and temperature is given. A method for adapting the data to
the case of oblique shocks is included. |
| 1127. |
Analytical and experimental
investigation of temperature recovery factors for fully developed flow
of air in a tube (Sep 1958) by W. F. Weiland, W. H.
Lowdermilk, R. G. Deissler [36 pages; 1.3 MB] |
|
Abstract: An analysis was made for predicting
temperature recovery factors for fully developed flow in a tube. Most
of the attention was confined to turbulent flow. Some qualitative
results were obtained for laminar flow by setting the eddy diffusivity
in the equation for turbulent flow equal to zero and using the
incompressible parabolic velocity profile for laminar flow. For zero
Mach number the laminar flow results were exact. Radial variation of
properties was neglected in most of the calculations. The effect of
wall temperature gradient along the tube was negligible for turbulent
flow below Mach numbers of 0.9 and 0.98 for Reynolds numbers of 20,000
and 390,000, respectively. For laminar flow the effect became important
at much lower Mach numbers. Recovery factors were obtained
experimentally for a range of Reynolds number from 630 to 30,000.
Additional previously unpublished data are presented for Reynolds
numbers up to 650,000. The results indicate that in the turbulent flow
region the recovery factor is approximately independent of Reynolds
numbers up to 650,000. In the transition region for Reynolds numbers
between 2000 and 3000 the recovery factor is reduced abruptly to a
value lower than that obtained for the turbulent flow region. |
| 1128. |
Preliminary heat-transfer
studies on two bodies of revolution at angle of attack at a Mach number
of 3.12 (Sep 1958) by Norman Sands, John R. Jack [30
pages; 1.5 MB] |
|
Abstract: Local rates of heat transfer were
obtained for a cone-cylinder model and a parabolic-nosed-cylinder model
at a Mach number of 3.12 and angles of attack up to 18 degrees. Data
were obtained for cooled surfaces at unit Reynolds numbers of 0.36 and
0.65 million per inch based on free-stream conditions. Zero angle of
attack data are included for comparison. |
| 1129. |
Comparison of shock-expansion
theory with experiment for the lift, drag, and pitching-moment
characteristics of two wing-body combinations at M=5.0 (Sep
1958) by Savin, Raymond C [14 pages; 0.5 MB] |
|
Abstract: Lift, drag and pitching-moment
coefficients for two wing-body combinations were determined from tests
at a Mach number of 5.0 and angles of attack up to 9 degrees. The test
models consisted of small thin wings mounted on a body composed of a
fineness-ratio-3 ogival nose and a fineness-ratio-2 cylindrical
afterbody. The wings were symmetrically mounted on the cylindrical
portion of the body and had triangular and trapezoidal plan forms. The
results of these tests are compared with results obtained by a
relatively simple application of the generalized shock-expansion method
in combination with the T' method of evaluating the skin-friction drag
coefficients. Good agreement between theory and experiment is obtained
for the total drag coefficients over the test angle-of-attack range.
Theory and experiment are also found to be in good agreement for the
lift and pitching-moment coefficients at the lower angles of attack. At
the higher angles of attack, the theoretically determined coefficients
are somewhat higher than those obtained experimentally. |
| 1130. |
Effects of nose angle and Mach
number on transition on cones at supersonic speeds (Sep 1958)
by K. R. Czarnecki, Mary W. Jackson [18 pages; 0.7 MB] |
|
Abstract: An investigation has been made to
determine the transition characteristics of a group of smooth,
sharp-nosed cones varying from 10 degrees to sixty degrees in included
apex angle over a Mach number range from 1.61 to 2.20 and a range of
Reynolds number per foot from about 1.5 x 10 to the 6th power to 8 x 10
to the 6th power. Increasing the cone angle is shown to decrease
slightly the transition Reynolds number, whereas the effects of changes
of Mach number and unit Reynolds number are negligible. When transition
occurred within 15 to 20 percent of the model length from the base
there was a dropoff in transition Reynolds number. (author) |
| 1131. |
Effect of advance ratio on
flight performance of a modified supersonic propeller (Sep
1958) by Jerome B. Hammack, Thomas C. O'Bryan [21 pages; 0.6 MB] |
|
Abstract: Results are presented of a flight
investigation to determine the aerodynamic characteristics of a
supersonic propeller modified by the incorporation of higher than
optimum advance angles. The propeller was designed for a forward Mach
number of 0.95, an advance ratio of 3.2, and a power coefficient of
0.42. The efficiency of the propeller is approximately 79 percent at a
Mach number of 0.95. At lower Mach numbers the efficiency is higher,
being about 85 percent at a Mach number of 0.75. The departure from
optimum angle of advance has a small effect for the advance ratios
investigated. |
| 1132. |
Use of the Kernel function in a
three-dimensional flutter analysis with application to a flutter-tested
delta-wing model (SEP 1958) by Donald S. Woolston, John
L. Sewall [43 pages; 1.4 MB] |
|
Abstract: The development and the numerical
application are presented of a Rayleigh-Ritz, or modal, type of flutter
analysis which takes into account three-dimensional structural and
aerodynamic behavior. The flutter mode is approximated by a series of
natural-vibration modes, and the aerodynamic forces corresponding to
these modes are derived from subsonic lifting-surface theory, according
to the kernel -function approach, for a finite wing oscillating in
compressible flow. The application is made to a delta semispan wing
with a leading-edge sweep angle of 45 degrees which fluttered at a Mach
number of 0.85. Results of flutter calculations show that, for this
case, when the first three or four natural-vibration modes are used to
approximate the flutter mode, converged solutions for the flutter speed
are obtained that are about 5 percent less than the experimental value.
Theoretical flutter-speed boundaries were located for a range of
densities and Mach numbers including those of the experimental-flutter
condition. Further application of the analysis to study the effects of
variation in certain structural properties showed that the converged
flutter speeds were more sensitive to variations in the natural
frequencies than to either variations in mass or to the inclusion of
generalized-mass coupling terms whose existence is due to the use of
experimental natural mode shapes. |
| 1133. |
Measurements of aerodynamic
forces and moments at subsonic speeds on a simplified T-tail
oscillating in yaw about the fin midchord (SEP 1958) by
Sherman A. Clevenson, Sumner A. Leadbetter [21 pages; 0.8 MB] |
|
Abstract: Results are presented of some
experimental measurements of aerodynamic forces and moments acting on a
simplified T-tail configuration which is oscillating in yaw about an
axis through the midchord of the vertical fin. Coefficients which
define rolling moment of the horizontal stabilizer alone and rolling
moment, yawing moment, and side force of the complete T-tail are shown.
In the investigation the range of reduced-frequency parameter was from
0.09 to 0.56, the Mach number range was from 0.13 to 0.50, and the
Reynolds number range was from 0.90 X 10-to-the-sixth to 8.21 X 10-to
-the-sixth. Coefficients for the steady case (reduced-frequency
parameter of zero) were calculated for the forces and moments and good
agreement was indicated for all except the horizontal-stabilizer
rolling-moment coefficient which was found to be of greater magnitude
than was indicated by the steady-state results. Some further
comparisons were made of the side-force and yawing moment on the
complete T-tail with results obtained from a previous investigation for
a configuration consisting of a tip tank mounted on a plan form similar
to the T-tail fin alone and were found to be compatible. |
| 1134. |
Free-flight investigation to
determine the drag of flat- and vee- windshield canopies on a parabolic
fuselage with and without transonic indentation between Mach numbers of
0.75 and 1.35 (SEP 1958) by Walter L. Kouyoumjian,
Sherwood Hoffman [35 pages; 1.7 MB] |
|
Abstract: A free-flight investigation was
conducted between Mach numbers of 0.75 and 1.35 to determine the
effects on model total drag and pressure drag of (a) canopy location
(along a parabolic body of revolution), (b) canopy windshield shape,
(c) canopy fineness ratio, and (d) transonic-area-rule indentation. |
| 1135. |
Flight measurements of the
vibratory stresses on a propeller designed for an advance ratio of 4.0
and a Mach number of 0.82 (SEP 1958) by Thomas C.
O'Bryan [15 pages; 0.6 MB] |
|
Abstract: Results are presented of
vibratory-stress measurements obtained in flight on a propeller
designed for an advance ratio of 4.0 and a forward Mach number of 0.82.
|
| 1136. |
Some measurements of
aerodynamic forces and moments at subsonic speeds on a rectangular wing
of aspect ratio 2 oscillating about the midchord (May 1958)
by Edward Widmayer, Jr., Sherman A. Clevenson, Sumner A. Leadbetter [46
pages; 1.3 MB] |
|
Abstract: Some measurements were made of the
aerodynamic forces and moments acting on a rectangular wing of aspect
ratio 2 which was oscillated about the midchord. These measurements
were made at four frequencies (31, 43, 54, and 62 cps) over a range of
Mach number from 0.15 to 0.81, a range of reduced frequency from 0.15
to 1.32 and a range of Reynolds number from 0.60 to 10 (sup 6) to 9.21
X 10 (sup 6). |
| 1182. |
Effect of target-type thrust
reverser on transonic aerodynamic characteristics of a single-engine
fighter model (January 13, 1958) by Swihert, John M [44
pages; 1.1 MB] |
|
Abstract: A brief investigation of a target-type
thrust reverser on a single-engine fighter model has been conducted in
the Langley 16-foot transonic tunnel at Mach numbers from 0.20 to
1.05.At Mach numbers of 0.80, 0.92, and 1.05, a hydrogen peroxide
turbojet-engine simulator was operated with the thrust reverser
extended. The angle of attack was varied from 0 degrees to 5 degrees at
these Mach numbers. The Reynolds number of the free stream, based on
the mean aerodynamic chord, was about 5 x 10(6). It was estimated that
reversed jet operations separated the model boundary-layer flow over
the upper surface of the horizontal tail and upper part of the
afterbody. This resulted in a positive pitch increment due to reversed
jet operation. Jet-on operation also tended to stabilize the severe
lateral oscillations which occurred with the reverser extended and the
jet off. It appeared that these jet-off oscillations were the result of
an alternating separation and reattachment of the flow on the rearmost
portions of the fuselage afterbody. |
| 1183. |
Effectiveness of various
protective coverings on magnesium fins at Mach number 2.0 and
stagnation temperatures up to 3,600 R (January 09, 1958)
by Bland, William M , Jr [49 pages; 1.2 MB] |
|
Abstract: No Abstract Available |
| 1184. |
Free-flight roll performance of
a steady-flow jet-spoiler control on an 80 degree delta-wing missile
between Mach numbers of 0.6 and 1.8 (January 24, 1958)
by Schult, Eugene D [38 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 1185. |
Wind-tunnel investigation of
the effects of lip geometry on drag and pressure recovery of a
normal-shock nose inlet on a body of revolution at Mach numbers of 1.41
and 1.81 (February 03, 1958) by Robins, A Warner [99
pages; 2 MB] |
|
Abstract: No Abstract Available |
| 1186. |
Some effects of fin
leading-edge shape on aerodynamic heating at Mach number 2.0 at a
stagnation temperature of about 2,600 R (January 09, 1958)
by Bland, William M , Jr [16 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 1187. |
Preliminary investigation of
graphite, silicon carbide, and several polymer-glass-cloth laminates in
a Mach number 2 air jet at stagnation temperatures of 3,000 F and 4,000
F (January 09, 1958) by Casey, Francis W , Jr Hopko,
Russell N [20 pages; 0.5 MB] |
|
Abstract: No Abstract Available |
| 1188. |
Experimental and theoretical
determination of forces and moments on a store and on a store-pylon
combination mounted on a 45 degree swept-wing-fuselage configuration at
a Mach number of 1.61 (January 30, 1958) by Morris,
Odell A Carlson, Harry W Geier, Douglas J [124 pages; 9.9 MB] |
|
Abstract: An investigation of store-pylon forces
and moments has been conducted in the Langley 4- by 4-foot supersonic
pressure tunnel at a Mach number of 1.61. Separate forces and moments
were measured simultaneously on a store and on a store-pylon
combination for a number of pylon-mounted store locations below the
wing of a 45 degrees swept-wing-fuselage combination. Tests were made
through an angle-of-attack range of -4 degrees to 12 degrees and an
angle-of-sideslip range of -12 degrees to 12 degrees. The basic model
configuration, which was almost identical to the model used in NACA RM
L55A13a, simulated a heavy-bomber-type airplane with a large ogive
cylinder store. The results of the investigation indicate that the most
important source of store-pylon side forces is the pylon itself. When
immersed in a strong sidewash field, the pylon can assume a large load
and also produce a large incremental load upon the store. Both tend to
increase rapidly with increasing angle of attack or angle of sideslip.
Location of the store-pylon combination in a sidewash field of strong
intensity may also result in powerful secondary effects on the normal
force and axial force of the store. The large unstable pitching moments
obtained for the sweptforward store-pylon installations at moderate
angles of attack indicate that release of an unfinned store from a
forward store location could be hazardous. Tests with two stores
mounted on the same wing panel show that the presence of the inboard
store-pylon combination causes significant decreases in the outboard
store and store-pylon side forces produced by angle of attack. The
theory as used here provides a useful estimation of the
angle-of-attack-induced store or store-pylon side force. However, the
side force is underestimated at the inboard wing positions and is
overestimated at the outboard positions. |
| 1189. |
Free-flight aerodynamic-heating
data at Mach numbers up to 10.9 on a flat-faced cylinder
(January 13, 1958) by Bland, William M , Jr Swanson, Andrew G
Kolenkiewicz, Roland [48 pages; 2.4 MB] |
|
Abstract: No Abstract Available |
| 1190. |
Measurements of aerodynamic
heat transfer in turbulent separated regions at a Mach number of 1.8
(February 20, 1958) by Garland, Benjamine J Hall, James R [18
pages; 0.4 MB] |
|
Abstract: No Abstract Available |
| 1191. |
An investigation at Mach
numbers 1.94 and 2.41 of jet effects upon the longitudinal and
directional stability of a general aircraft configuration employing
wing-tip-mounted nacelles (March 04, 1958) by Clark,
Frank L Edwards, Clyde L W [71 pages; 2 MB] |
|
Abstract: No Abstract Available |
| 1192. |
Heat transfer measured on a
flat-face cylinder-flare configuration in free flight at Mach numbers
from 1.6 to 2.7 (February 03, 1958) by Lee, Dorothy B
Swanson, Andrew G [35 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 1193. |
Effects of nose and afterbody
modifications on aerodynamic characteristics of a body with and without
a vertical tail at a Mach number of 2.01 (April 15, 1958)
by Foster, Gerald V [87 pages; 10.8 MB] |
|
Abstract: No Abstract Available |
| 1194. |
Heat transfer to 0 degree and
75 degree swept blunt leading edges in free flight at Mach numbers from
1.90 to 3.07 (March 24, 1958) by O'Neal, Robert L Bond,
Aleck C [39 pages; 2 MB] |
|
Abstract: No Abstract Available |
| 1195. |
Wind-tunnel investigation at a
Mach number of 2.01 of the aerodynamic characteristics in combined
angles of attack and sideslip of several hypersonic missile
configurations with various canard controls (March 10, 1958)
by Robinson, Ross B [35 pages; 2.5 MB] |
|
Abstract: No Abstract Available |
| 1196. |
Jet effects on the base drag of
a cylindrical afterbody with extended nozzles (April 15, 1958)
by Nelson, William J Scott, William R [43 pages; 1.2 MB] |
|
Abstract: A wind-tunnel investigation to
determine the effects of both single and twin jets on base drag of a
cylindrical body has been conducted at Mach numbers from 0.6 to 1.4.
The plane of the jet exit was varied with respect to that of the
afterbody. Jet total-pressure ratio ranged up to 20. Significant
improvements in base drag were obtained by extending the plane of the
jet exits beyond the afterbody base and by venting the base cavity to
the external stream. |
| 1197. |
Heat transfer for Mach numbers
up to 2.2 and pressure distributions for Mach numbers up to 4.7 from
flight investigations of a flat-face-cone and a hemisphere-cone
(May 08, 1958) by Speegle, Katherine C Chauvin, Leo T Heberlig,
Jack C [61 pages; 2.6 MB] |
|
Abstract: No Abstract Available |
| 1198. |
Heat transfer and pressure
measurement on a 5-inch hemispherical concave nose at a Mach number of
2.0 (July 17, 1958) by Markley, J Thomas [22 pages; 1.1
MB] |
|
Abstract: No Abstract Available |
| 1199. |
Investigation of control
effectiveness and stability characteristics of a model of a low-wing
missile with interdigitated tail surfaces at Mach numbers of 2.29,
2.97, and 3.51 (July 14, 1958) by Presnell, John G , Jr
[29 pages; 2 MB] |
|
Abstract: No Abstract Available |
| 1200. |
Tests of aerodynamically heated
multiweb wing structures in a free jet at Mach number 2 and
aluminum-alloy model of 40-inch chord with 0.125-inch-thick skin
(June 23, 1958) by Griffith, George E Miltonberger, Georgene H
[40 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 1201. |
Effect of aerodynamic heating
on the flutter of a rectangular wing at a Mach number of 2
(June 23, 1958) by Runyan, Harry L Jones, Nan H [18 pages; 0.5
MB] |
|
Abstract: No Abstract Available |
| 1202. |
Longitudinal and lateral
aerodynamic characteristics at combined angles of attack and sideslip
of a generalized missile model having a rectangular wing at a Mach
number of 4.08 (August 14, 1958) by Smith, Fred M
Ulmann, Edwards F Dunning, Robert W [93 pages; 2.2 MB] |
|
Abstract: No Abstract Available |
| 1203. |
Measurements of the buffeting
loads on the wing and horizontal tail of a 1/4-scale model of the X-1E
airplane (September 17, 1958) by Rainey, A Gerald Igoe,
William B [37 pages; 0.9 MB] |
|
Abstract: The buffeting loads acting on the wing
and horizontal tail of a 1/4-scale model of the X-1E airplane have been
measured in the Langley 16-foot transonic tunnel in the Mach number
range from 0.40 to 0.90. When the buffeting loads were reduced to a
nondimensional aerodynamic coefficient of buffeting intensity, it was
found that the maximum buffeting intensity of the horizontal tail was
about twice as large as that of the wing. Comparison of power spectra
of buffeting loads acting on the horizontal tail of the airplaneand of
the model indicated that the model horizontal tail, which was of
conventional force-test-model design, responded in an entirely
different mode than did the airplane.This result implied that if
quantitative extrapolation of model data to flight conditions were
desired a dynamically scaled model of the rearward portion of the
fuselage and empennage would be required. A study of the sources of
horizontal-tail buffeting of the model indicated that the wing wake
contributed a large part of the total buffeting load. At one condition
it was found that removal of the wing wake would reduce the buffeting
loads on the horizontal tail to about one-third of the original value. |
| 1204. |
Free-flight investigation of
aerodynamic heat transfer to a simulated glide-rocket shape at Mach
numbers up to 10 (September 10, 1958) by Swanson, Andrew
[50 pages; 2.4 MB] |
|
Abstract: Heat-transfer measurements were made on
a simulated glide-rocket shape in free flight at Mach numbers up to 10
and free-stream Reynolds numbers of 2 x 10(6) based on distance along
surface from apex and 3 x 10(4)based on nominal leading-edge diameter.
The model simulated the bottom of a 75 degree delta wing at 8 degrees
angle of attack. The data indicated that for the test conditions a
modified three-dimensional stagnation-point theory will predict to
reasonable engineering accurary the heating on a highly swept wing
leading edge, the heating being reduced by sweep by the 3/2 power of
the cosine of the sweep angle. The data also indicate that laminar
heating rates over the windward surface of a highly swept flat glider
wing at moderate angles of attack can be predicted with reasonable
engineering accuracy by flat-plate theory using wedge local flow
conditions and basing Reynolds numbers on lengths from the wing leading
edge parallel to the surface center line. |
| 1205. |
Heat transfer measured in free
flight on a slightly blunted 25 degree cone-cylinder-flare
configuration at Mach numbers up to 9.89 (September 26, 1958)
by Lee, Dorothy B Rumsey, Charles B Bond, Aleck C [63 pages; 3.9 MB] |
|
Abstract: No Abstract Available |
| 1206. |
Stability investigation of a
blunt cone and a blunt cylinder with a square base at Mach numbers from
0.64 to 2.14 (September 17, 1958) by Coltrane, Lucille C
[33 pages; 1.4 MB] |
|
Abstract: No Abstract Available |
| 1207. |
Rocket-model investigation to
determine the lift and pitching effectiveness of small pulse rockets
exhausted from the fuselage over the surface of an adjacent wing at
Mach numbers from 0.9 to 1.8 (September 30, 1958) by
Martz, C William [24 pages; 0.6 MB] |
|
Abstract: No Abstract Available |
| 1208. |
Theory and experiments on
supersonic air-to-air ejectors (1958) by Fabri, J.,
Paulon, J. [31 pages; 1 MB] |
|
Abstract: A comparison of experiment with theory
is made for air ejectors having cylindrical mixing sections and
operating under conditions of supersonic primary flow and either mixed
or supersonic regimes of mixing. The effect on ejector performance of
such parameters as mixer length and cross section, terminating
diffuser, primary Mach number, and primary nozzle position is presented
in terms of mass flow and pressure ratio. |
| 1209. |
The interaction of a reflected
shock wave with the boundary layer in a shock tube (1958)
by [131 pages; 4.3 MB] |
|
Abstract: By theoretical analysis the existence
of several different types of interaction in different ranges of
initial shock Mach number is predicted. This analysis is verified
experimentally. The most compicated interaction is studied in detail,
and a model is proposed. The features of the phenomenon are analyzed,
based on this model, and are checked experimentally. |
| 1210. |
Turbulent skin friction at high
mach numbers and reynolds numbers (1958) by Matting,
Fred W., Chapman, Dean R. [8 pages; 0.3 MB] |
|
Abstract: (abstract not available) |
| 1211. |
Static stability and control of
canard configurations at Mach numbers from 0.70 to 2.22 :longitudinal
characteristics of a triangular wing and canard (1958)
by Boyd, John W., Peterson, Victor L. [34 pages; 1.5 MB] |
|
Abstract: The canard configuration investigated
at Mach numbers from 0.70 to 2.22 consisted of a triangular wing and
triangular movable canard, both of aspect ratio 2.0, a low-aspect-ratio
vertical tail, and a fineness ratio 12.5 Sears-Haack body. |
| 1212. |
A wind-tunnel investigation of
several wingless missile configurations at supersonic speeds
(1958) by Reese, David E., Jr. [53 pages; 2.3 MB] |
|
Abstract: A wind-tunnel investigation of several
wingless missile configurations was made. Lift, drag, and
pitching-moment coefficients were measured on a series of models at
Mach numbers of 2.44 and 3.35 and on one model from 1.76 to 5.05. |
| 1213. |
Transonic investigation of
yawed wings of aspect ratios 3 and 6 with a Sears-Haack body and with
symmetrical and asymmetrical bodies indented for a Mach number of 1.20
(1958) by Holdaway, George H., Hatfield, Elaine W. [113 pages;
3.3 MB] |
|
Abstract: Two yawed wings, each with an average
sweep of about 40 degrees, were investigated with various bodies and
the results were compared with existing data for similar models with
sweptback wings. |
| 1214. |
The static longitudinal
characteristics of a twisted and cambered 45 degree sweptback wing at
Mach numbers up to 0.96 (1958) by Sammonds, Robert I.,
Reynolds, Robert M. [28 pages; 1.7 MB] |
|
Abstract: The wing-body combination tested had a
wing of aspect ratio 3 incorporating twist and a distributed type of
camber. |
|
|
|
Total number of pages: 47014
pages |
|
Total size: 1923.2 MB |
|
|
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