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UNCLASSIFIED AD 658 995 FINAL ENGINEERING REPORT ON WIND TUNNEL TEST STUDY Collins Radio Company Cedar Rapids, Iowa June 1958 Processed for . . . DEFENSE DOCUMENTATION CENTER DEFENSE SUPPLY AGENCY FOR FEDERAL SCIENTIFIC AND TECHNICAL INFORMATION U. S. DEPARTMENT OF COMMERCE / NATIONAL BUREAU OF STANDARDS / INSTITUTE FOR APPLIED TECHNOLOGY UNCLASSIFIED
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UNCLASSIFIED
AD 658 995
FINAL ENGINEERING REPORT ON WIND TUNNEL TEST STUDY
Collins Radio Company Cedar Rapids, Iowa
June 1958
Processed for . . .
DEFENSE DOCUMENTATION CENTER DEFENSE SUPPLY AGENCY
FOR FEDERAL SCIENTIFIC AND TECHNICAL INFORMATION
U. S. DEPARTMENT OF COMMERCE / NATIONAL BUREAU OF STANDARDS / INSTITUTE FOR APPLIED TECHNOLOGY
UNCLASSIFIED
.
UBRARY COPY Ä ^.
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Report No. CER-826 Part I Copy No. 17
PART I
FiMAL ENGINEERING REPORT
ON
WIND TUNNEL TEST STUDY ^)
1 June 1958
DA-44-177-TC-448 P.O. DEE-10017
Covering Period from 25 January to 1 April 1958
DECLASSIFIED
§E6tAS§lFieATIQ^ DATE £!$&*p.JS.&2...
(COLLINS EMPLOYEE • DftTE) /
NOTICE: This document contains information affecting the national defense of the United States, within the meaning of the Espionage Laws, Title 18, U.S.C., Sections 793 and 794, the transmission or revelation of which in any manner to an unauthorized person is prohibited by law.
COLLINS RADIO CO, io JAN 1960
GROUP IV DOWNGRADED AT ?, YEAR INTER-
VALS; DECLASSIFTED AFTER 12 YEAR?.
DOD DIR 5200.10
8118
CEDAR RAPIDS. IOWäJ
C Reproduced by the
CLEARINGHOUSE for Federal Scientific & Technical Information Springfield Va. 22151
'iLINS AERONAUTICAL RESEARCH LABO^Toiv A. M. Lipplsch, Director ._Cedar Rapids, Iowa
C l
this document has been approved for public release end solo; its distribution is unlimited.
OCT 41967 1^
UN JNFIDENTIAbv liiu*—^»
Si. M ^
■ - ■ ■
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A PUBLICATION OF
THE AERONAUTICAL RESEARCH LABORATORY
COLLINS RADIO COMPANY
Cedar Rapids, Iowa
PiMUd In th« Unilfd StaUs of Amwiea
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FINAL ENGINEERING REPORT ON
WIND TUNNEL TEST STUDY
1. INTRODUCTION.
This is the first part of the final repor. which includes information pertaining to the con-
struction and testing of the CARV*. This wind tunnel test model was constructed and tested
according to the specifications set forth by the Chrysler Corporation (Chrysler PO DEE-
100017 under prime contract DA-44-177-TC-448).
The initial program for testing the CARV* called for: (1) modification of existing wind
tunnel facilities to allow testing of the 1/10 scale model, (2) testing the model to determine
its aerodynamic characteristics in pitch to 30°, in yaw to 90°, and in roll to 30°; these tests
to be performed with forward speeds between 0 and 50 miles per hour, and (3) testing the
model to determine the control vane effectiveness.
As is the case in testing most unorthodox configurations, the scope of the program was
altered considerably during the testing as the results began to indicate fields of more
specialized interest.
Modification of the existing wind tunnel facilities was undertaken to allow for six compon-
ent measurements in pitch and In roll attitudes.
The first part of the test program as performed up until March 31, 1958 covered the
following investigations:
(a) Vane effectiveness in hovering.
(b) Forward flight conditions for tandem configuration.
(c) Forward flight conditions for abreast configuration.
(d) Finding a method to reduce the positive pitching moment.
(e) Miscellaneous.
♦Chrysler Aerial Research Vehicle
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The following professional and technical personnel worked on this project during the period
covered by this report.
Dr. A. M. Lippisch
Mr. E. A. Sielaff
Mr. J. C. Brown
Mr. R. N. Pugh
Mr. R. F. Colton
Mr. L. H. Connover
Mr. L. L. Miner
Mr. W. C. Green
Mr. A. G. Johnson
Mr. M. V. Hoppenworth
Mr. J. A. Felthous
Director of the Collins Aeronautical Research Laboratory
Senior Engineer
Senior Assistant Engineer
Senior Assistant Engineer
Junior Assistant Engineer
Aeronautical Engineering Assistant
Laboratory Assistant
Laboratory Assistant
Laboratory Assistant
Laboratory Assistant
Draftsman
Draftsman
0
Mr. E. A. Sieh
2. IDENTIFICATION OF SYMBOLS.
The symbols used in this report may be Identified by referring to the following list. Figures
1 and 2 will also aid in Identifying the symbols.
0
Symbol
C =-£- C pu2S
C =^- D 2« pu S
C --i- L Ä 2S pu
1 /0u2Sd
Dimensions
lb
Definition
Side force
Side force coefficient in hovering
Drag coefficient in hovering
Lift coefficient in hovering
Rolling moment coefficient in hovering
0
2
, ,._ ,._ w.
O
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Symbol Dimensions Definition
C = ^ m pu2Sd
Pitching moment coefficient in hovering
c - N' n u2Sd
Yawing moment coefficient in hovering
D lb Drag
K --£- KD-%S Drag coefficient for forward flight
K - L Lift coefficient for forward flight
KN-qoS Normal force coefficient for forward flight
K -L' Rolling moment coefficient for forward flight
Pitching moment coefficient for forward flight
L lb Lift
L' lb ft Rolling moment
M' lb ft Pitching moment
N ) Normal force
N' 3 ft Yawing moment
S ft2 Area of the two shrouds com- bined measured in the plane of the propellers
e
ft
ft
RPM
% lb/ft„and kg/m2
Diameter of the shrouds measured in the plane of the propeller
Distance of the normal force from the C.G. (+ forward)
Revolutions per minute of the propellers
Dynamic pressure of the air from wind tunnel (measured)
_ . -ßONTlDENTIAU
Symbol
u
v.
0
a
8
p
Dimensions
lb/ft2 and kg/m2
ft
ft/s
ft/s
ft/s
ft/s
degrees
degrees
degrees
degrees
degrees
lb s2/ft4
degrees
radians/s
Definition
Dynamic pressure of the slip stream of the propeller
Radial distance from propeller axis to blade section
Propeller tip speed
Tunnel speed reduced to stand- ard sea level condition
Axial air speed through propeller
Tangental velocity of blade section of propeller
Angle of pitch of the model
Effective angle of attack of a section of the propeller
Deflection angle of the vanes
Angle between the chord of the propeller blade section and the plane of rotation
Angle of blade setting of the propeller at the hub
Air density during the test
Effective pitch angle of the propeller
Angular velocity of the propeller
Subscripts:
A, B, C, D in connection with S indicate pitch vanes.
1, 2, 3, 4 in connection with S indicate roll vanes.
3. TEST APPARATUS.
Some construction was necessary in order to adapt the existing wind tunned facilities to
the testing program outlined for the model. A general description rf the testing system and
the construction phase of the test model can be found in IDR-449-1 dated February IS, 1958.
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The following Information is In addition to the information supplied in the previous
report.
3.1 WIND TUNNEL DESCRIPTION.'
The wind tunnel utilized for testing the model Is an open throat, open circuit tunnel which
has a throat diameter of 36 Inches. The range of wind velocities available for testing purposes
is 1.5 feet/second to 40 feet/second. The maximum error encountered in measuring the
wind velocities was ±1%. The test model was supported so that its center of gravity was
located on the center line of the tunnel and 32 inches from the plane of the mouth of the
tunnel. The model was suspended 60 inches from the floor.
3.2 MODEL MODIFICATIONS.
In addition to the facilities described in IDR-499-1 for adjusting the vanes to various
angles, more recent modifications made it possible to vary the position of the vanes in a
vertical direction (see figure 3). The vanes can either be set at the original position,
(position 1), at the intermediate position (position 2) which places the leading edge of the
ry vane In the plane of the shroud exit, or at the lower position (position 3), which places the
leading edge of the vane one chord length below the shroud exit (see figure 4). The vane
chord lengths were Increased by one chord length for certain phases of the testing program.
3.3 SCALE SYSTEM.
Figures 5 and 6 show the scale arrangement that was used to obtain the force and moment
data. The scales are harvard trip balance, double beam types and have a capacity of two
kilograms. The rated sensitivity of the scales is 0.1 of a gram and the error in measuring
accuracy does not exceed 0.1%. For the tests performed, stops were added to the scales
in order to limit their range of travel. The locations of each scale and their respective
functions are shown in figure 7.
A special frame was used to adjust the pitch angle of the model. This frame was mounted
on the scale platform (see figure 8).
e>'»K,iaiM)3tCM*r;
ISuLfsüoiI Iti
OONnDENTIAl«' . . . ,
L 3.4 MEASURING APPARATUS.
CONEIJDE;
o The equipment used to monitor the operating voltage and current for the model motors,
the motor rpm rates, and the wind velocities below the shrouds and in the wind tunned are
shown in figures 9 and 10. Magnetic pickup devices were installed on one motor shaft to
facilitate the recording of the rpm rate. The outputs from these devices were fed to a
spectrum analyzer which is capable of displaying information between the range of 2000 rpm
and 110,000 rpm. The accuracy of the equipment was within 0.2%.
A regulated, variable power supply unit furnished the necessary operating voltages for
the two DC motors used in the test model. A vernier control was also employed to aid in
regulating the voltage output of the power source. Power was measured with a voltmeter
and ammeter, which provided accurate power data to within ±0.5%.
Pressure surveys for the areas directly below the model shrouds were accomplished with
the rake attachment shown in figure 11, which was attached to the shroud during the test
runs. The rake assembly was designed so that each tube represents an equal area of flow
from the shroud. The tubes of the rake were attached to the multiple manometer shown in r\
figure 11. The accuracy of the readings obtained was within ±0.25 millimeters of HgO.
A Prandtl tube, used in conjunction with a Betz manometer, provided a means of checking
the wind velocities in the tunnel. The manometer has a range of 0 to 800 millimeters of H-O.
The minimum accuracy obtained was within ±0.1%.
4. TEST PROCEDURE.
The following is a description of the procedure used to obtain data during a performance
test. These data were used to plot the curves of dimensionless coefficients. Although a
particular test run is considered here for purposes of illustration, the methods described
apply generally to the entire testing program.
The first part of the procedure outlined below (steps 1 through 8) was necessary to esta-
blish a performance point. The second part of the procedure was required to establish the
IdK ÖKD dK \ jtability derivatives I-^r , —TT I —Sg I around this point. äk
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Step 1. With the wind tunnel turned off the lift scales were checked for zero values for
\J the range of pitch angles to be covered during the test. These values were measured and
recorded so that the proper weights could be added to the scales during the test runs. This
procedure ensured that the measured lift values were net readings.
Step 2. The model was then set at the -20° pitch angle with the proper zero weights on
the lift scales.
Step 3. The DC generator for the tunnel motor was started, warmed up, and switched to
the tunnel blower motor.
Step 4. The approximate propeller speed was determined (computations based on pre-
vious work) which would deliver the required 1600 gms. lift. The model motors were then
energized.
At this point it might be well to consider the reasoning involved in setting up the forward
flight performance test program. It was felt that the only point of practical significance was
the condition of steady horizontal flight. This eordition is obtained when the drag on the
model is equalized by the thrust produced by the propellers and inlets. This condition can be
stated as follows: Drag - Thrust = Drag measured =0. In addition to this, the lift was
maintained at a constant value (1600 gms.) for each of the performance data points. This
magnitude of lift was selected so that it would be as large as possible and still not exceed the
range of available power in the model motor for all conditions of forward and hovering flight.
The remaining steps are required to produce the conditions outlined above.
Step 5. After completing step 4 the two drag scales were unblocked and the tunnel wind
speed was gradually increased until both drag scales were balanced simultaneously. If a
slight amount of yawing moment existed, the weights on the drag scales were adjusted so that
the force added on one scale was subtracted from the other scale.
Step 6. When the exact wind velocity was determined at the condition D = 0, the drag
scales were blocked and the four remaining scales were read in order. While these scales
were being read, the wind velocity and prop rpm rate were held constant. Also, at this time,
the wind velocity and the power input to the motors was recorded.
TOfFIF
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Step 7. After the lift and side force measurements were made, the power to the model
motors was turned off and the exact amount of lift was computed. The direction and approxi- ^
mate magnitude of the change required to bring the lift to 1600 gm was then determined.
Step 8. By an iteration process of this kind, it was usually possible to attain the re-
quired lift (±0.5%) after no more than four tries.
Step 9, Once the tunnel wind speed and the prop speed for the steady horizontal flight
condition at this angle was determined the six components were measured for 2.5° and 5°
angle increments above and below the original -20° pitch angle. The two parameters men-
tioned above (rpm, tunnel r.peed) were held constant for these additional readings.
Step 10. The raw data obt-ined from the tests were recorded on standard data sheets.
These data were analyzed, interpreted and finally plotted on graph paper.
5. TEST RESULTS
The material in this section summarizes the tests performed up to March 31, 1958.
These tests covered Investigations in the following areas:
(a) Vane effectiveness in hovering for pitch, roll and yaw.
(b) Forward flight conditions for tandem configurations which involved:
(1) Determining the zero drag conditions, and the change in lift, drag and pitching
moments due to varying the angle of attack and the forward speed.
(2) Vane effectiveness in pitch.
(c) Forward flight conditions for abreast configurations involving:
(1) Determing the zero drag conditions and the change in .lift, drag and pitching
moments due to varying the angle of attack and the forward speed.
(2) Vane effectiveness in pitch.
(d) Finding a method to reduce the positive pitching moment.
(e) Miscellaneous.
Each of the areas of investigation listed above are discussed in detail in the following
material. Section 6 of this report includes graphs which were drawn from the data obtained
from the various test runs.
8 . • > ..CONFIDEHflTÄL &
" dDKEiDENJIA&
5.1 DISCUSSION OF TEST RESULTS.
(a) Vane effectiveness in hovering for pitch, roll, and yaw. The model configuration for
these test runs was as specified for the original design and included four roll vanes
and two pitch vanes in each shroud. The propeller setting was «>% = 30°. The model
was suspended in the hovering position and the tests were preformed at prop speeds
of 4000, 5500, and 7000 rpm. The tunnel wind speed was zero.
(1) Yaw effectiveness.
Figures 16 and 17 show the lift and yawing moment coefficients for 4000, 5500,
and 7000 rpm as a function of the vane deflection. Vanes 1 and 2 in the front
shroud and 3 and 4 in the rear shroud were deflected by equal angles in opposite
directions. The pitch vanes A, B, C, and D were in the neutral positions.
(2) Roll effectiveness.
Figures 18, 19, and 20 show the lift, side force and rolling moment coefficients
for 4000, 5500, and 7000 rpm a« a function of the vane deflection. All roll vanes
(1, 2, 3, and 4) were deflected by the same angle and in the same direction. The
vanes A, B, C, and D were in the neutral positions. The effect of deflecting the
roll vanes on one side only is shown in figure 21. The sets 1 and 3 were deflected,
and all other vanes were in neutral positions. These test runs were performed
at prop speeds of 7000 rpm only and for positive deflections.
(3) Pitch effectiveness.
Figures 22, 23, and 24 represent the lift, drag and pitching moment coefficients
for 4000, 5500, and.7000 rpm. The pitch vanes A, B, C, and D are deflected by
the same angle and in the same direction. All roll vanes are in a neutral position.
Since the pitching moment coefficients show considerable scattering and indicate
no vane effectiveness for negative deflections, more tests were performed at
prop speeds of 7000 rpm only to see if any improvement could be achieved.
Figures 25 and 26 show the coefficient curves which resulted from deflecting the
pitch vanes in the rear shroud in opposite directions. All other vanes are in
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0 neutral positions. Blocking of the rear shroud by deflecting all vanes in the
shroud inward resulted in a pitching moment coefficient curve as shown in figure
27. The graphs in figures 28 and 29 represent some of the significant coefficients
obfei .ed by deflecting the pitch vanes in the rear or front bhroud only. Finally
the vanes in the rear shroud were lowered to position 3 (see figure 4) and the
double pitch vanes were installed (see figure 12). The measurements for this
configuration were taken several times and the pitching moment coefficients
obtained are plotted in figure 30. The curves which should be identical scatter
widely and show no consistency.
(4) General investigations concerning vane effectiveness in hovering. There were a
large number of runs which were not included in the data point numbering system
simply becuase these runs were either pressure measurements or spot checks
of various configurations designed to provide one particular result.
In the initial phases of this work it was felt that the control vanes were ineffec-
tive because the mass flow in the region of the vanes was too small. Conse-
quently, the effort in the first phase was directed toward increasing the velocity
near the wall relative to the central part of the stream. This velocity increase
can be accomplished in at least two ways. One method is to make suitable inlet
modifications. Another is to redistribute the pitch of the prop blades to provide
for more work at the tips.
The inlet modifications were tried first. A summary of the configurations tested
is shown in figures 31 and 32 and the curves corresponding to the dynamic pres-
sure distributions for those configurations for which data were recorded are
shown in figures 33, 34, and 35. A detail drawing of the total head rake used to
obtain this pressure information appears in figure 35A. The second method of
accomplishing the desired velocity distribution is to redistribute the pitch along
the prop blades. This "prop twisting" was done after an optimum inlet configura-
tion had been decided upon, with one exception, that being that the prop in the
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rear shroud had been "straightened out" according to the pitch distribution no. 1
C"^ (see figure 36) previous to the inlet tests and was used during all of these runs.
All of the rest of the prop distribution work was done with the final ring form
(see figure 32) above the inlet.
The prop pitch analysis was done as described in the following text. In all of the
work the induced angle of attack increment was neglected. Figure 15 along with
the following definitions will serve to identify the terms in the equations used in
the ensuing discussion.
•& = angle between plane of rotation and chord line of the blade section.
ä = effective angle of attack of this blade section.
4> = angle between plane of rotation and relative wind at prop blade section (effective pitch angle).
27mr vT = ü,r="60-
va=v/2^ fa
tan<A=va/vT= ^
Since all of the test runs were performed at a constant prop speed (n = 7000 rpm)
the above expression can be simplified to the form:
By measuring q1 with the total head rake at various stations it is then only a
matter of using the above equation to determine the angle of the relative wind at
that station.
Assuming that it is desired to obtain a uniform velocity distribution with the same
thrust that is provided with the actual distribution, a curve can be drawn which
represents the 0 distribution for uniform flow.
Having established the desired incurve, and also curves for the measured values of
\^) «9-and <£, the following procedure can be used to determine the required ^distribution.
■
mvi IVsXiHrl
At any station
a actual ~ meas " ^meas' \J
This expression allows us to determine the actual angle of attack. With this
Information •$■, . , Is determined from the equation.
desired ~ ^desired actual
This value of ^9-Is the angle at which the prop pitch must assume at that particular
station in order that the desired velocity distribution may be approached. This
process of iteration Is continued until the desired velocity distribution is obtained.
Normally this process allows the particular velocity distribution to be approached
quite rapidly (see figures 36, 37, and 38).
The graph in figure 36 shows the curves that were obtained from the initial con-
dition,. The graph in figure 37 represents the distribution after the first twisting
procedure and the graph shown in figure 38 represents the final distributions.
This prop configuration was used for both props for all tests from no. 170 on. /ps
Figure 39 is a plot of the C versus 8 curve after the modifications mentioned
above had been performed. The graph illustrates how A for the front prop was
changed to provide for C^Xo at S = o.
(b) Forward flight conditions for tandem configurations,
(1) Determining the zero drag conditions.
The zero drag conditions were determined for the original shroud and vane con-
figurations with all of the vanes in the neutral position. Figure 40 shows the
lift and pitching moment coefficient as a function of the pitch angle, the drag com-
ponent being zero. Figures 41 and 42 show these same functions with the rpm of
the props held constant (4000 and 7000). The graph in figure 43 represents the
tunnel velocity (reduced to standard sea level conditions) as a function of the
pitch angle with the rpm rate of the props at 4000 and 7000 and with the drag equal
to zero. f}
12 CÖNFIDENTDOL^.
ISS PsyLWööii
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The travel of the center of pressure for the different steady horizontal flight
conditions at various pitch angles is shown in the graph of figure 44.
The change of the lift, drag and pitching moment coefficients with variations of
the pitch angle and forward speed was determined in the following manner:
Keeping the lift component constant (1.600 kg.), the zero drag conditions were
determined for four different angles of pitch (Ö = -10°, -20°, -30° and -37.5°), for
which the corresponding tunnel velocities were v = 10.6, 17.4, 23.8 and 28.4
ft/s. Then the angle of pitch was varied ±5° in 2.5° increments from the original
angle, with the tunnel wind velocity and the prop rpm rate being held constant.
The resu'ts of these tests are shown in figure 45. Additional test runs were
made holding the angle of pitch and the prop rpm rate constant and varying the
tunnel speed by two increments above and below the zero drag velocity. The re-
results of these tests are shown in figures 46 and 47. The vanes were placed in
position 2 (see figure 4).
(2) Vane effectiveness in pitch.
The effect of deflecting the pitch vanes in the rear shroud only is shown in figure
The vanes were in position 3 (see figure 4) and double pitch vanes had been
lied. It can be seen that deflecting the vanes in this case did not produce any
positive results as far as effectiveness was concerned. Measurements were also
taken with only the pitch vanes in the front shroud deflected. The double pitch
vanes remained in position 3. In this case the measurements were taken only
with the vanes deflected in the negative direction. Figure 49 shows that no satis-
factory results were obtained.
A series of tests were performed with all the double pitch vanes deflected. No
change in effectiveness was apparent during these tests (see figure 50). Vertical
fins were added to the Inner vanes of sets A and D, but did not have any substan-
tial effect (see figures 51 and 13). The curve of figure 52 show the coefficients
13
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with all of the vanes in position 2 (see figure 4) and deflected by equal angles in
the same direction. O
In order to investigate the effect of the size of the vanes on vane effectiveness,
the chord lengths of vanes C and D were doubled (see figure 14). The data shown
in figure 53 were obtained with the vanes placed in position 2. The correspond-
ing curves for vanes of normal chord length are also shown. The effect of dou-
bling the chord length of all of the vanes is shown in figure 54.
(c) Forward flight conditions for abreast configuration.
(1) Determining the zero drag conditions, change in lift, drag and pitching moments
due to varying the angle of attack and the forward speed.
The zero drag conditions for the tests performed in this phase of the work were
determined using the original shroud configuration. The double pitch vanes were
installed and all vanes were placed in position 2 (see figure 4) and had zero de-
flection. Figure 55 shows the lift and pitching moment coefficient as a function
of the pitch angle, the drag being zero. Figure 56 represents the forward speed,
reduced to standard sea level conditions, as a function of the pitch angle at con-
stant lift. Figure 57 shows e/d as a function of the pitch angle at zero drag.
The change in the lift, drag and pitching moment coefficients for small varia-
tions in the pitch angle and forward speed was determined in the same way that
these coefficients were found for the tandem arrangement. The results of the
tests are shown in figures 58, 59, and 60.
(2) Vane effectiveness in pitch.
The vane effectiveness in pitch was determined by first deflecting the vane sets
1, 2, 3, and 4 (see figure 2), which act in the abreast arrangement as pitch
vanes, by the same angle and in the same direction. The pitch angle used was
-20*. The results of these tests are shown in figure 61.
For additional testing, the results of which are shown in figure 62, the vanes f}
were initially deflected outward by 10*, i.e., the undeflected condition was
14
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iä *>, S I »■ Ö iia it
S = S = +10° and 8^ = 5 = -10°. From this position the vanes were deflected
Cj by equal amounts in the same direction. No improvement was obtained,
(d) Finding a method to reduce the positive pitching moment.
In order to eliminate the nose up pitching moment the blade setting of the front prop
was varied from 33.25° to 25.25° at the hub in increments of 2°. The results obtained
with the prop set at these various angles are shown in figure 63. The pitching mo-
ment was reduced to zero by reducing the prop angle 4.6°. Figure 63 also shows the
corresponding increase in the rolling moment for the reductions in the pitch
angle.
The second phase of this work was directed toward finding the cause of the largf.
nose up pitching moment and determining what could be added in the way of lift re-
ducers to diminish this tendency. The configurations tested are shown in figure 64
and the results of the tests are shown in figure 65. All of the tests from 1 through
13 were performed with a prop speed of 7000 rpm and a pitch angle of -25°. The
series of tests for configuration 14 were performed with a prop speed of 5710 rpm
and a pitch angle of -20s. The shrouds were in the abreast position for these
tests.
The rest of the time spent in the pitching moment investigations was devoted to taking
pressure measurements at the shroud inlets to see if the pressure existing at the lips
was sufficient to account for the strong nose up pitching moment in forward flight.
The pressure distributions obtained from these tests are shown in figures 66, 67, and
68. Table 5-1 lists the significant data obtained.
15
T'
TABLE 5-1. PITCfflNG MOMENT PRESSURE DATA
Test No.
Shroud Arrangement
Pitch Angle
Prop rpm Tunnelq
N shroud Ntotal
<e/d)shroud
Ntotal
1 tandem 0 5712 0 .30
2 tandem -20* 5712 1.55^ .75
3 abreast 0 5900 0 .34
4 abreast -20° 3900 1.55^; M2
.18 .70
o
(e) Miscellaneous.
(1) The results of the power measurement tests are plotted in figure 69. During
these tests, the lift was maintained at a constant value of 1600 gms. and the
drag component was held at a value of zero. The upper curve in figure 69 re-
presents the data obtained with the tandem configuration. Each angle on the
curve represents a different angle of pitch.
(2) In order to determine the effect of upward speeds on the lift characteristics in
hovering, the model was suspeV with the inlets of the shroud facing the tunnel
mouth. The force parallel to ti of the shrouds (lift) was measured for
different wind speeds. Figure 70 shows the lift as a function of the upward speed.
(3) Another series of tests was run in order to check whether the position of the
tunnel axis had any effect on the measurements. Data were taken for the same
vane deflections, prop speed, tunnel speed and pitch angle. However, the model
was adjusted 3, 6, and 9 inches above and below the tunnel axis for vane deflec-
tions of 0s and 30s (see figure 71). In can be seen that the slope of the curves is
practically zero for the central position of the model, which indicates that this
position yields the correct test data.
O
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6. ILLUSTRATIONS.
O
TANDEM CONFIGURATION
o
+6t ,+64
+6l,-'-61
ABREAST CONFir ^TION
o
Figure 1. Symbol Identification (Forward Flight Condition)
CONFIDENTIAI/
IMSSlFJiD 17
r. t?« n \r
Figure 2. Symbol Identification (Hovering Flight C s)
0
FRONT REAR
^4
'"""••••ttn
+
+
+ oo +
:^
o
0
18
o
r "™ i 1
O
;
P v;3
I.H -»■N'.X
\
A s
o
Figure 3. Closeup View of Model Shroud Showing Attachments for Adjusting Vanes in Vertical Direction
(Vanes in Position #3)
19
■' ■ ■■
POSITION POSITION 2 POSITION 3 0
Figure 4. Vane Positions
1 p.aamj Bfi »M ^•T-iifV ifVI mum
irf\x. :—t VN, . XMllV
.*.><
"^
> tlöl si-
Figure 5. Wind Tunnel Scale System (Looking to Rear)
0
0
20
li
o
o
o
""••C
—a " i ^ ■ rv--
3 iLiiÄ^,
/ /
^-;. ;/: i:
\\
iZ /•-, •''' £~''L>.' "1 « i
i '■>
I v r r*& *
i "^
Figure 6. Wind Tunnel Scale System (Looking Forward)
21
o
FRÖNT LIFT SCALE,LEFT FRONT LIFT SCALE, RIGHT REAR LIFT SCALE DRAG SCALE. LEFT
DRAG* SCALE,RIGHT DRAG COUNTERBALANCE
SIDE FORCE SCALE SIDE FORCE COUNTERBALANCE
FRAME FOR ADJUSTING ANGLE OF ATTACK
0
Figure 7. Wind Tunnel Scale System 0
22 'BRB
lllyLildy r r
_
o
TO LIFT SCALE REAR
TO LIFT SCALE LEFT FRONT
TO LIFT SCALE RIGHT FRONT
O
TO MODEL REAR
TO MODEL LEFT FRONT
TO MODEL RIGHT FRONT
Figure 8. Pitch Adjustment
O
flDENTKL,
JlHiLHdöirirJl 53
F r —v T« i *•■—^A« - -.—^-^i^^. ,,r——-
I'.ii*'.-,' .:f^ Si fcr--.-
v ,. i
>".,„>
\ -•vr
:v i ^V^ *- A. -I 1 » \ C-Tr.-~V- T
1 ■■n p:3i
It»; .: r \
IP7**?.
'i fcä. ES > Ö"
1.
1 I . ■ I .ft n ■ /
1 ^
I ^ J '
uoöl. ^ s^^ r i
7
■ „^ -•> -^ 1:
: ("T
, -v ■ i:
i •
_,y
: t.
■**». ■ l^l' , Ifc». ,^ *■■.■.■<!. <>. 4
Figure 9. Wind Tunnel Testing Facilities
o
0
0
24
O
o
o
Figure 10. Measuring Equipment
«.CQNFIDENTIAI» 25
-:
0
...i... i . «. .1. J u i ■ l.i.L
^
// y/
t^C '/•-C/
-f
">!'
1
ii 1 D.w^:
0
Figure 11. Total Head Rake and Multiple Manometer
0
26
Ml*l k^K \^r^ i'-
o K tamzatOB*
o Figure 12. Double Pitch Vanes
Figure 13. Fins Above Inner Vanes of Set A and D
O _i
Figure 14. Vanes with Double Chord Length
27
WMm&mlp
^.„w,—Hmi
o
o
Figure 15. Prop Section with Symbols used in Related Calculations
O
28
._. ... __
ÜiiyU^biLu CONFIDENTIAL^
I03CL
♦ 20
♦ 10
♦30°
-30°
♦20
-20
♦ 10
-10
-10
♦ 10
-20 Ma -30°
♦ 20 8j84 ♦30°
Figure 16. C, vs Vane Deflection for Yaw in Hovering
29
.»q^-^^-WKf'WWW
w«>hi.>h^4^.>.«i.^tX..»w—■■ i- ill i ii'iniMiriimlfclilMfiMlBiiMMfcaM^
U- . . C^r^LJS.-Z.&L
io5.cB
+ 190
m-
+ 100
+ 50
-SO
-100
-150
1— CHRYSLER AERIAL TEST VEHICLE WIND TUNNEL TEST •»o » 0 YAW
8] 82 +30° +20« +10° 0 -10° -20O -30° -30° -200 -10° 0 + 5° +10° +15° +20° +250+30a
8384
o
0
Figure 17. C vs Vsme Deflection for Yaw in Hovering 0
o
n KJi " r:» C«\
o
+ 10 ♦ 20 ♦30° 8, >. ». »,
o
Figure 18. CL vs Vane Deflection for Roll in Hovering
•CONFIDENTIAL 31
limraimif CONHDENTIAL
32 D^NTIAL.
0
0
Figure 19. C vs Vane Deflection for Roll in Hovering
0
yi^ÄÄsi/^
IO»C| ii i + 60
mmm. CHRYSLER AERIAL TEST VEHICLE
WIND TUNNEL TEST
4 40
4 20
-20
-40
-60
420° 430°
8| 82 83 *4
o
Figure 20. Cj vs Vane Deflection for Roll in Hovering
CONnDENTiAE 33
' V'i Ii w 8 liA n H H q B 1 M
\ 1 r i r
34
ill PTIXXITIK II
O
O
8, 8, ♦30«
Figure 21. C, vs Vane Deflection for Roll in Hovering. Roll Vanes Deflection on One Side Only
o
o
o
♦20° tSO0
'* 8B 8C 80
o
Figure 22. C^ vs Vane Deflection for Pitch in Hovering
liyillddiritL 35
im PSSMllu li^LÖNFlbfiNÄL
36
O
IO'-CB
♦ 0.5
-0.5
0
Figure 23. Cp vs Vane Deflection for Pitch in Hovering
0
MI u mmammmtm ,^1 « w» yi
/'U <J ^-/
io»cm
♦ 80
♦ GO
♦ 40
♦ 20
Figure 24. Cm vs Vane Deflection for Pitch In Hovering
IOC,
+?o
mmmmmmmmm CHRYSLER AERIAL TEST VEHICLE WIND TUNNEL TEST vo »0 PITCH
■TTTTT
+10
-30° -20 -10 0 ♦ 10 ♦ 20
♦ 30» ♦ 20 ♦ 10 0 -10 - 20
♦ 30°
30" 8,,
Figure 26. CL VS Vane Deflection with Rear Shroud Pitch Vanes Deflected in Opposite Directions
„ _ „CGNFIDENTIA-Ia
iflel«! II \ \ i if-ili 37
Imvl M ^\j^ir li '
lo'Cn, IO3C0
♦ 80
♦ 60
♦ 40
♦ 20 ♦O.Z
•0.2
-0.4
-30" -20 -10 0 ♦ 10 ♦20 ♦30° 8c
♦ 30° ♦ 20 ♦ 10 0 -10 -20 -30° 8o
Figure 26. CQ and Cm vs Vane Deflection with Rear Shroud Pitch Vanes Deflected in Opposite Directions
o
o
o 38
liäilSSliiEii eONFlDENTIAXk
o
l05Cm
♦ 80
♦ 60
♦ 40
+ 20
8C 83 +30°
»0 84 -30°
O
Figure 27. Cm vs Vane Deflection with Vanes in Rear Shroud Deflected Inward
39
r« ^ p !-'To
40 afnsmiF 16! r 'S B "'•■■! E 3 1 \i y L'u. uiw oil ts-^
o
o
Figure 28. C^ vs Vane Deflection with Pitch Vanes Deflected in Rear or Front Shroud Only
0
C 6;
o
lO5^ IO3C0
+80 +0.8
+ 60 +0.6
+ 40 +0.4
+ 20 +0.2
O -20 •0.2
-0.4
-0.6
-0.8
4tiWi.-litii)t!^ffli|-Plii^iilltt- CHRYSLER AERIAL TEST VEHICLE
WIND TUNNEL TEST
Sc SQ-SO0
8A 8B-30o +30° 8C 80
o
Figure 29. CD and Cm vs Vane Deflection with Pitch Vanes Deflected in Front or Rear Shroud Only
CONFIDENTIAI» 41
V V- it W D 4 m 0
o
o
Figure 30. Cm vs Vane Deflection with Vanes in Rear Shroud at Position 3 and Double Pitch Vanes Installed
O
42
Um** r-SSi fir 5
TESTS 1-5
O
^
PLAIN SHROOD
^ ^
o
o
TESTS 6-10
10
/P
SAME AS ;^0.9 BUT NOT FILLED
rr*
Figure 31. Shroud Inlet Modification
CONFfDENTlAL
lflirM*$friffi 43
fi'. H y a n a
N ö>w©mtjEAr>
O
TEST NO. 11-16
12 13 14 15 16
SAME AS NO. 10
BUT 5 IN. CUT FROM
TOP OF RING
/^
SAMEASNO.il
BUT ■j IN. ADDITIONAL
CUT FROM TOP OF RING
/P
SAME AS NO. 12 BUT SMALLER
RADIUS OF CURVATURE
SAME AS N0.I2 BUT LARGER
RADIUS OF CURVATURE
rr* rr>
FINAL FORM OF RING FOR BEST
RESULTS
O rr
Figure 32. Shroud Inlet Modification (Cont)
O
44 CONFIDENTIAL
m i
't&xAmwi&iS
o
o
o
2 3 4 INCHES OUT FROM «.
-+- I 2 3 4 5 POSITION OF NUMB^HED TOTAL HEAD TUBES ON RAKE
Figure 33. Dynamic Pressure Distribution Curves Obtained from Configurations in Figure 31 with Prop Pitch Distribution #1
45
[ GONFIDENTIAt.
iä tSi>
JL _l_
3 4
INCHES OUT FROM (. J-
I 2 3 4 S
POSITION OF NUMBERED TOTAL HEAD TUBES ON RAKE
Figure 34. Dynamic Pressure Distribution Curves Obtained from Configurations in Figure 31 with Original Prop Pitch Distributions
O
O
o 46
W B CONFIDENTIAD
18' I Hi
tff
:;;
Tin it::
illi -r.r
| M
if—I— CHRYSLER AERIAL TEST VEHICLE
WIND TUNNEL TEST
■ t ;
wri
ÖStffi
iSaStSIiruMt^i 4» s ^Skm
2 3 INCHES OUT FROM %_
I I II I 2 3 4 POSITION OR NUMBERED TOTAL HEAD TUBES ON RAKE
^fOYNAMIC PRESSURE DISTRIBUTION ACROSS SHROUD FOR VARIOUS IN- LET CONFIGURATIONS-USING THE ORIGINAL PROP WITH ALTERED PITCH DISTRIBUTION NO. I PROP SPEED s 7000 RPM TUNNEL SPEEDs o
sfifflssum
i 5
Figure 35. Dynamic Pressure Distribution Curves Obtained from Configurations in Figure 32 with Prop Pitch Distribution #1
„ n (V, eONF-IDENTIAl, „ 47
^' ^bimöEkiui^ ü ii L
o
o
Figure 35A. Detail of Total Head Rake
O
48
o
o
o
2 3 PROPELLER STATIONS
1 1 8 3
1 4
1 6
TOTAL HD TUBE LOCATIONS
Figure 36. First Propeller Pitch Modification Curves
n f. f*. '> CÖNFIDENTIAIu,
muH fic?c?§F:pcn 49
büVOGNFIDENTlAli11 L
LJ U a o a
u -i o z <
2 3 PROPELLER STATIONS
O
O
2 3 TOTAL HO TUBE LOCATIONS
Figure 37. Second Propeller Pitch Modification Curves o 50 QONFIDENTIÄL
mm R^CIElf:
2 3
PROPELLER STATIONS 1— 1
2 3 TOTAL HD TUBE LOCATIONS
Figure 38. Final Propeller Pitch Modification Curves
CONFIDENTIAI/,
fM I « I m > ^ R« i I« I i
51
IsViSd WfilfClE IILM « i^!i • 5 !• J . TV ,.\i Hit
52 M • I'
0
5 Äffl rn
Ott
••■80
1-60
+ 40
1-20
-20
-40
-6 0
•80
•100
m m rtt
m f;-t-<-
1 ;il :-T.:
ffi tttt
ÄJ
Effi
MrrHimilKFimHHHi;::;iFffl]mH^iHfi}i^: -t—^1 PUDVCI CP ACOIA1 TCCT W P U I r* I C +H-r
^n^' m 443
frP
is
1
tw
Stfff fffl
CHRYSLER AERIAL TEST VEHICLE WIND TUNNEL TEST
•'o = 0 PITCH n « 7000 RPM
P SB
m m
m
vrt TUJ
m m im
U mm mi
tt 1
i ffi ttt
Sf Bit
SB iffi Sf
FEET
p^;
O NO. 170-173 O NO. 178-161
S m B i-h-ttT
m 9i
I£;ffil3^ :ttH
^
■J
H-H -H-rr
3j&+&r^SB: ± m± ttF,
FRONT PROP PITCH 31.25° AT HUB
mttiffl
fr1
ttc?
nrr
±;4
rjfej^n JIT ?! I: rHf'StjltftlSl:
^HTaE
r fa
rr:: —
i,1!!
-Hi4
t*
K
ipc
ffi
:5l
;-iT
tttTTjT
+ 10° + 20° + 30°
o
SC 8D
Figure 39. Variation of C with S for Deflection
of Pitch Vanes in Rear Shroud Only
O
CGNHDENTIAL
O
Figure 40. K, and K vs 6 for Forward Flight in Tandem
Configuration and Zero Drag
53
CONFIDENTIAL^ ^
o
F
2.0
1.5
1.0
0.5
n
Km 25
20
15
10
t p-fi
Wi-'imM
CHRYSLER AERIAL TEST VEHICLE WIND TUNNEL TEST TANDEM
mm. P
1 3^
o
Figure 41. K, and K vs Ö for Forward Flight in Tandem
Configuration, Zero Drag and Constant RPM (n = 4000)
o 54
V F :;
■ i Li s v 1 fi ii
2\J ^20NPibENTIÄt? ^ ^ ^
o
♦•2.0
+ 1.5
O ♦ 1.0
♦ 0.5
♦ 20
♦ 15
♦ 10
♦ 5
o
Figure 42. K, and K vs Ö for Forward Flight in Tandem
Configuration, Zero Drag and Constant RPM (n = 7000)
•v 2a^QW^WH'fir$m 55
ft 11" ' K:
y ^^ (CONripENTIALi ^^^
56
MHHMHH
Figure 43. v vs 6 for Forward Flight in Tandem Configuration, Zero Drag and Constant RPM (n = 4000 and 7000)
^CONFIDENTIAL I i m a ß rA \. \ ■ I > I»
O
O
o
o
ti:!C':cbNFIDiNTlA^ rP^ I
^n
o
o
Figure 44. e/d vs Ö for Forward Flight in Tandem Configuration, Zero Drag and Constant RPM
(n = 4000 and 7000)
CÜNFlDENTIAIca,
UPI $Qonr§rn 57
V 0. f" •: w
3 n laOJIFIÖENTl^
58
ili#P|igii 24
Ko«Km + 3.0
+2.0
+ 1.0
0 - 0
- -1.0
J-Z.O
Figure 45. Change of KT , K- and K Due to Variation of 6
for Forward Flight in Tandem Configuration
. CONFIDENTTÄL
O
o
o
o
r r t f f- "- •^ ■"* T •■ •
I 1; l-i t wfoKFl^ENTTAL1;' t ^ vl W-' wa ^ •* V. ■"■■■'..•■'"•■ ti li i»ai
-
o
+40
+ 30
+20 +2.0
+ 10 +1.0
K0 + 1
0 0 0
-I
-2
O
Figure 46. Change of KL, K_. and K Due to Variation of v
for Forward Flight in Tandem Configuration
CONFIDENTIAL 59
{ i I, i r "'■
11 $ ICCNFIDEKTI^i^ *■«« B
60 •CONFIDENTIAL
O
+40
+30
+20 +ao
+10 +2 +1.0
+ 1
0 0 0
-2
CHRYSLER AETRIAL TEST VEHICLE WIND TUNNEL TEST TANDEM
.10 .15
rrr .25 .30 35 ,40 2
I I I W6 .20 .25 .30
O
Figure 47. Change of K , K_ and K due to Variation of v
for Forward Flight in Tandem Configuration
O
mvrr-f! ft^Cü'CIErt «;J u\.-.-.,■ i-.-j ü u v.- O' 6 a J ifea ta£/
0
o
^T •'m *L ''o 7
0.3
10 I
0.2
s as
i o.i
0 0 0 0
rtr IM
9.= 3.65 K8/m« n « 7000 RPM ö = -25° V, = 25.1 fl/»
TANDEM
-Hl 1 CHRYSLER AERIAL TEST VEHICLE
WIND TUNNEL TEST
H;:
ffi
tu- rn
Figure 48. K., K-, K and e/d vs Vane Deflection for Forward Flight in Tanden Configuration. Double Pitch Vanes in Position 3 in Rear Shroud Deflected
o
m
K« KL KQ -J
2 0.2
0 0 0 0
MHiil 9» = 3.33 Kg/m2
n = 7000 RPM 0r — 25° TANDEM
u\ lii; ]]\] iill iliy ♦4+" Til"'- ■ ■ i i ■—- ■■ ''■■■■
Em Sjj^isiiiiMtrKfitjlSH
CHRYSLER AERIAL TEST VEHICLE WIND TUNNEL TEST
hi
iSljlMM EM ;.!-:: U::
+ 20 8A. 8B
Figure 49. K,, K , K and e/d vs Vane Deflection for Forward Flight in Tandem
Configuration. Double Pitch Vanes in Position 3 in Front Shroud Deflected
lj5ßßmFIDENTIAt
; :ö^bL^dd.lrll:ll 61
.^^niiTin cvi ;r.n f:; ;"-IE\H s alia
o
o
8A8B8C80
Figure 50. K,, KD, K and e/d vs Vane Deflection for Forward Flight in Tandem
Configuration. Double Pitch Vanes in Position 3 Deflected in Both Shrouds
O
62
ga.^,CGNFIDENTTAbi|r y
o
o
Figure 51. KT and K vs Vane Deflection for Forward Flight in Tandem Configura-
tion. Double Pitch Vanes Deflected in Both Shrouds. Fins Added.
O
GONFIDENTIAL
Mm iwinsrH 63
I ■ r\i •A V. r-v
y ['COß'FMwiJlU l Ü
o
IK L Q j
♦ 10+10
o
Figure 52. K-, KD, K and e/d vs Vane Deflection for Forward Flight in Tandem
Configuration. All Vanes in Position 2 Deflected in Both Shrouds
O
64
o
r-"
y s^ip^ö^ ä s
o
KLK0Km
10 1.0
S 0.5
0 0 0
O
Figure 53. K-, K--, and K vs Vane Deflection for Forward Flight in Tandem
Configuration .Comparison Between Normal and Double Chord Length Vanes Deflected in Rear Shroud
65
n
66
h •..•' .i-' .1 vv,: v' ti ;u 1 !i; d
O
♦10 +1.0
♦ 6 ♦0.5
+ 20 8A 8B^8D+30o
Figure 54. KL, KD, K vs Vane Deflection for Forward Flight in Tandem Configuration.
Double Chord Length Vanes Deflected in Both Shrouds.
O
O
o
yenNFIDEkTlAI^ii [f* | f
o 3.0 30
2.0
1.0 10
0 0
O
Figure 55. Ky and K vs 9 for Forward Flight in Abreast Configuration
at Zero Drag
^C.WIQENTIAJL Imp* /■ Cti ^ fl r-Sp m
67
eONFIDENTIAIU-
o
c
Figure 56. v vsfl for Forward Flight in Abreast Configuration at Zero Drag and Constant Lift
%
♦ 0.2
► 0.1
i T^rft^Tnti^^l;:::!':::!::::!;:!7!^^ CHRYSLER AERIAL TEST VEHICLE WIND TUNNEL TEST A0REAST
o Figure 57. e/d vs 0 for Forward Flight in Abreast Configuration at Zero Drag
68
is.^'^'cörrnDEMTiAur? ^> m
o
o
o
-2.0
Figure 58. Change of KL, K.. and K Due to Variation of 6 for Forward Flight
in Abreast Configuration
OWFIDENTTAln.
AV & b vj 'Ls, i.rc <■-.»■ sv;" s 2 ST« "J
69
än.y* S ! ISS ^ölE'iltiii)
70
o
^L KD Km
+ 30
+20 +2.0
+ 3
+ 10 +2+ |,0
+1
0 0 0
-3
o
Figure 59. Change of K,, K- and K Due to Variation of v for Forward Flight
in Abreast Configuration
o
o xm
o
50
440
+30 + 3.0
420 + 2.0
+ 3
+10 +2 +1.0
+ 1
-I
-2
o Figure 60. Change of Kj, K_ andK Due to Variation of v for Forward Flight
in Abreast Configuration
71
.. r« r; ff\ r; '.^ v
''bciStrlDJiNTIAL ■
o
KL KO Km
+ 10 +1.0
+ 2.0
+ S +0.5
+ 1.0
-0.5
-1.0
O
Figure 61. K, , Kn and K vs Vane Deflection for Forward Flight in Abreast Configuration.
Pitch Vanes Deflected in Both Shrouds
o 72
'■
yj. ■' ' i ". ! ■ '. :l >, -.1;, '; « t' j fe ,: .CONHpENJlA^; £a g f li
o
o
♦ IJO *\0.0
»m
■»■2.0
f0.8 *8J0
•0.5
*I.O
j^m^iii» CHRYSLER AERIAL TEST VEHICLE
WIND TUNNEL TEST
-10° -5 0 + 5 + 10 + 15 +20 +25 + 30° -SO«» -25 -20 -15 -10 -5 0 + 5 + 10°
8, 8, 8o 8/
Figure 62. K^, K^ and Km vs Vane Deflection for Forward Flight in Abreast Configuration. Pitch Vanes in Both Shrouds Deflected Outward by 10*
O
73
• P. \. ' ■) ' j iGDimnEJi'nAp '■- * tu
"5" ^m ^L ^O
«0.9
«10 «1.0 ■tfi
♦0.2 «2
♦0 «OB
♦0.1 ♦!
CHRYSLER AERIAL TEST VEHICLE WIND TUNNEL TEST
q^» 3.3S KO/m,2
n = 7000 RPM fl«-250
TANDEM TTTT
-0.1-I
♦ o.i
♦ 0.05
-0.06
^o OF FRONT PROP
fflALL VANES AT ZERO DEF
1
llltllllllim.lllllHi.llMlllli.lll.nUTTTTTTT O
o
25.25° 27.25°
Figure 63. KL. KD, K^, e/d and Ki vs ^j, for Forward Flight in Tandem Configuration with all Vanes in Neutral Positions
o 74
mini ^'p.^r'V- i ILO h W
0
0
Na SKETCH OF CONFIGURATION COMMENTS ON CONFIGURATION NO. SKETCH OF CONFIGURATION COMMENTS ON CONFIGURATION
tj
CZD CZZ1
-I««
far
PLAIN JO WITH RINGS PATT "TER THE FIN IN PART NO. 1 NC SEDUCERS.
SAME AS NO! WITH FLAT SHEET LAYING ON TOP OF STRUTS IN SHROUD.
SAME AS NO.2 WITHOUT RINGS ABOVE INLET
PLAIN SHROUD WITHOUT RINGS INLET AND WITHOUT BAFFLE
PLAIN SHROUD WITHOUT RINGS BUT WITH THRUST REVERSER-OPERATION VERY ROUGH
PLAIN SHROUD WITH 1/2" CYLINDER PROJECTING ABOVE FRONT 1/4 OF SHROUD FLUSH WITH INSIDE WALL.
iAME AS NO. 3 WITH BAFFLE PERFORATED WITH 1/4" HOLES 1/2'APART
10
12
13
14
jpr
nzn i—p
[ZZ1.•,
1
17"-»I h m>
—IN IO
PLAIN SHROUD WITH FLAT BAFFLES ARRANGED AS SHOWN
FILLED IN SPACE BELOW FRONT LIP. COVERING FRONT HALF OF SHROUD
ATTEMPT TO STREAM- UNE REAR OF MODEL
FILLED IN SPACE BELOW REAR 1/2 OF REAR SHROUD UP
CONTOURED SPACE BELOW REAR 1/2 OF REAR SHROUD LIP
PLAIN SHROUD WITH I CYLINDER PROJECTING ABOVE FRONT 1/2 OF SHROUD FLUSH WITH INSIDE WALL
SHROUDS ABREAST.LARGE VANE ADDED ACROSS REAR AS SHOWN
a. a • 35* b. a • 45« e. a ■ 55» d. a • 20*
0 Figure 64. Pitching Moment Reducing Devices
CONFIDENTIAL 75
,
ii i\ \:J\ ■'<" H r--
UJi iiidu
in CD
Ii ill! ll i mm iiiHH Htt
TTTT ii CHRYSLER AERIAL TEST VEHICLE
WIND TUNNEL TEST TANDEM
Oiii mm ilii ID TUNNEL VEL. HD (q0) = 2.5 Kg/Mz
A TUNNEL VEL. HD (q0)=3.33 Kg/M2
D TUNNEL VEL. HD (q0)= 1.55 Kg/M2
6 8 l(
MOMENT (FT LBS)
0
o
Figure 65. Results of Pitching Moment Reducer Tests o 76 „ CONFIDENTIAL
H I !,.n yi:\;^^Kv::;;.^: j
o
o
/""S U Figure 66. Inlet Pressure Distribution on Front Shroud in Tandem
mbUWümtu 77
78
CHRYSLER AERIAL TEST VEHICLE' WIND TUNNEL TEST
Figure 67. Inlet Pressure Distribution on Rear Shroud in Tandem
^CONFIDENTIAL
■ ^J L^b ■ukninJ ;; s ^ (C ,8.1
O
o
o
0
0
0
ft I'vT'hM'^TTS'C'K'TTAt i- il '•- '< ri 5
CHRYSLER AERIAL TEST VEHICLE WIND TUNNEL TEST
-A Ä- FORWARD FLIGHT
Figure 68. Met Pressure Distribution on Left Shroud in Abreast
CONFIDENTIAL^ t;," n 79
' 1 -• rxo %'!v
' 'CON'FIDE. 'lAÜ " B fa '**
o
o
Figure 69. Power Ratio for Forward Flight in Tandem and in Abreast O
80 CONFIDENTIAL
f'. n :v l ij ,. ;■. .' ; , .. _<, f; '' h b ÖE
I; !:•■' h iv» ,1
0
0
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Figure 71. KL, K vs Distance of CG. of Model from Center Line of Tunnel
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