Comparison of measured and predicted indicated angles of attack near the fuselages of a triangular-wing wind-tunnel mode...

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Material Information

Title:
Comparison of measured and predicted indicated angles of attack near the fuselages of a triangular-wing wind-tunnel model and a swept-wing fighter airplane in flight
Series Title:
NACA RM
Physical Description:
13 p. : ill. ; 28 cm.
Language:
English
Creator:
McFadden, Norman M
McCloud, John L
James, Harry A
Ames Research Center
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

Subjects / Keywords:
Airplanes -- Wings, Triangular   ( lcsh )
Fighter planes -- Wings -- Testing   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract:
Abstract: Comparisons are presented of predicted values of flow angles and those measured using detectors close to the surface of the forward portion of fuselages of a large-scale, triangular-wing, wind-tunnel model at low speeds and a 35° swept-wing fighter airplane in flight. Effects of flap deflection on the upwash of the triangular-wing model are included.
Bibliography:
Includes bibliographic references (p. 7).
Statement of Responsibility:
by Norman M. McFadden, John L. McCloud, III, and Harry A. James.
General Note:
"Report date March 11, 1953."
General Note:
"Classification changed to unclassified Authority: NACA Research Abstract No. 97 Date: February 24, 1956."--stamped on cover

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003808552
oclc - 129843456
sobekcm - AA00006170_00001
System ID:
AA00006170:00001

Full Text

-A-\A z,..coPy344 J.
SECURITYv INFORMATION RM A53A15










RESEARCH MEMORANDUM


COMPARISON OF MEASURED AND PREDICTED INDICATED ANGLES
OF ATTACK NEAR THE FUSELAGES OF A TRIANGULAR-WING
WIND-TUNNEL MODEL AND A SWEPT-WING
FIGHTER AIRPLANE IN FLIGHT

By Norman M. McFadden, John L. McCloud, III,
and Harry A. James
Ames Aeronautical Laboratory
Moffett Field, Calif.
UN RSnY OF FLORIDA
DOC MENTS DEPARTMENT
120 ARSTON SCIENCE LIBRARY CLASSIFICATION CHASED TC UNCLASSIFIED
O. X117011
l ESV'_E.0 FL 32611-7.011 USA AUTHORITY NACA RESEARCH ABSRACT NO. .57
DATE: FERUART 24, 1956 I
CLASSIFIED DOCUMENT
This material contains formation affecting the National Defense of the United States within the meaning
of the espionage laws, Ttle 18, U.S.C., Secs. 79B and 794, the trawmlnslon or revelation of which in any
in to un uh.ored person la prohibited by law.

NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON
March 11, 1953











NACA RM A53A15


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


RESEARCH MEMORANDUM


COMPARISON OF MEASURED AND PREDICTED INDICATED ANGLES

OF ATTACK NEAR THE FUSELAGES OF A TRIANGULAR-WING

WIND-TUNNEL MODEL AND A SWEPT-WING

FIGHTER AIRPLANE IN FLIGHT

By Norman M. McFadden, John L. McCloud, III,
and Harry A. James


SUMMARY


Measurements of the local flow angles near the fuselages of a
triangular-wing wind-tunnel model and of an F-86A-5 airplane in flight
have been made by the use of air-flow detectors on the fuselages.
Comparison of these measured flow angles are made herein with predicted
flow angles. The methods used accurately estimated the change in upwash
due to flap deflection on the triangular-wing model. However, the
potential flow equations used to estimate the upwash due to the presence
of the fuselage consistently overestimated the effects of the fuselage
at the location of the detectors, which were very near the fuselage but
outside the boundary layer. At the detector location on the wind-tunnel
model (1.2 semispans forward of the quarter-chord line) the change in
upwash due to flap deflection was small and could be neglected for most
applications.


INTRODUCTION


Many cruise- and fire-control systems for aircraft and guidance
systems for missiles contain computers requiring signals proportional
to the true angle of attack. Reference 1 has shown that one feasible
method of obtaining the true angle of attack is by the use of detectors
which measured the local flow angle (i.e., the indicated angle of
attack) near the nose of the fuselage. It was found that for the test
airplane the true angle of attack was a linear function of the local
flow angle.


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There are two questions regarding this method of obtaining the
required angle-of-attack signals. One question concerns the possi-
bility of estimating the local flow angle accurately enough to make
flight calibration unnecessary, or, at least, accurately enough to
enable the choice of the location of the detector to be made with
confidence in the airplane design stage and, furthermore, to enable
the design of the required computers to be made without waiting for the
flight calibrations of the angle-of-attack system. The second question
concerns its use on triangular-wing aircraft where eleven deflection
might possibly influence the flow at the detector location to such an
extent that the calibration and computers required to reduce the local
flow angles to true angles of attack would have to use signals propor-
tional to flap deflection as well as angle of attack.

This report presents the results of low-speed wind-tunnel measure-
ments of local flow angles using a detector mounted on the forward
portion of the fuselage of a triangular-wing-fuselage model. These
test results and test results of reference 1 are compared with values
of indicated angle of attack predicted by the methods of reference 2.


NOTATION


c mean aerodynamic chord

CL lift coefficient, lift/1 pV2) S

S wing area, sq ft

V velocity, ft/sec

a angle of attack, deg

5f flap deflection, deg

E upwash angle, deg

p density, slugs/cu ft


Subscripts


b body

G geometric, referenced to tunnel center line


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I indicated

T true

w wing


APPARATUS AND TESTS


An aspect ratio 2 triangular-wing-fuselage model was tested in the
Ames 40- by 80-foot wind tunnel with a Specialties, Inc., Type J,
Airstream Direction Detector, used to indicate the local flow angle,
mounted on the fuselage 1.2 semispans forward of the wing quarter-chord
line. Dimensions of the model are given in figure 1.

The Airstream Direction Detector is a pressure actuated null
seeking device. The detector has a small cylindrical probe with two
lengthwise slots spaced 600 apart which provide differential pressure
to rotate the probe to seek the null or zero differential position which
is recorded by a potentiometer.

The local flow angles were measured at a tunnel dynamic pressure
of 30 pounds per square foot over an angle-of-attack range of 00 to 80
and with a flap deflection range of 15 Since the detector was
mounted well forward of the wing, the wind-tunnel-wall corrections at
the location of the detector were negligibly small and were not incor-
porated in the results.


THEORETICAL ESTIMATES OF THE LOCAL FLOW

Method


A method of estimating upwash in the extended wing-chord plane for
airplanes with swept wings is presented in reference 2. The method
assumes that the total upwash angle is the sum of the individual upwash
angles due to the presence of the wing, fuselage, and nacelle with the
fuselage and nacelle acting in the upwash field of the wing. Lifting-
surface theory of reference 3 is used to obtain the upwash due to the
wing, while the upwash angles due to the nacelles and fuselage are
obtained as in reference 4 from potential-flow equations, assuming the
fuselage to be an infinite cylinder and the nacelles to be semi-infinite
bodies of revolution. The method has been extended in reference 5 to
cover regions above and below the wing-chord plane.


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Application of the Method


The method of references 2 and 3 was used to obtain the upwash
effects due to the wing on the wind-tunnel model. However, since the
detector was rather near the nose of the fuselage, the semi-infinite
body theory of reference 4 was used to obtain the upwash due to the
influence of the fuselage rather than considering the fuselage to be an
infinite cylinder.

As in reference 2, the local flow angle is equal to the geometric
angle of attack plus the upwash due to the presence of the wing and
fuselage; that is,

aI = aG + Ew + Eb

which can be written
dew deb dEw L
aI= G + dC L + + d-CL

Then
a dEw dCL db dew dCL
L= 1 + + (1 + dC- (1)
1 + dCL daO da dCL
and
&LI dew dCL dEb d w dCL
+ (2)
56f dCL d6f do dCL dSf

Equation (1) is used to determine the slope of the curve of
indicated angle of attack versus true angle of attack. The variation
of upwash with lift coefficient, dew/dCL, is obtained from the lifting-
surface theory as applied in reference 2 (assuming an elliptic span
load); the lift curve slope, dCL/da, is obtained from experimental
results (fig. 2); and the variation of upwash with angle of attack,
deb/da, due to the presence of the fuselage, is obtained from potential
flow equations for a semi-infinite body of revolution (ref. 4).

The effect of flap deflection on the local flow angle is obtained
from equation (2). Again dew/dCL is obtained from lifting-surface
theory (ref. 2), assuming an elliptic span load; the change in lift
coefficient with flap deflection, dCL/d6f, is obtained from experimental
results (cross plot of fig. 2); and dEb/da is obtained from potential-
flow equations (ref. 4).

The method was applied to the F-86A-5 in essentially the same
manner as for the wind-tunnel model except that the span load and lift-
curve slope for the wing were obtained by use of reference 3.


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RESULTS AND DISCUSSION

Triangular-Wing Model


The angle of attack indicated by the detector on the wind-tunnel
model is shown as a function of the geometric angle of attack in
figure 3. For comparison, the predicted variation is also shown for
zero flap deflection. Although the agreement is not exact, it is
considered adequate for preliminary calculations. The zero shift of
about 0.2 is attributed to the dissymmetry of the fuselage just aft
of the location of the detector. (Since the shape of the fuselage
at a moderate distance behind the detector has little effect on the
calculated upwash, the fuselage was assumed to be a body of revo-
lution as shown in fig. 1 for the purpose of calculation.) As for the
discrepancy between the measured and predicted slopes of 0.10 (meas-
ured daI/daG = 1.73 and predicted = 1.83), most of the error is
thought to be in the predicted value since the estimated accuracy of
the measured slope is 0.02. The contribution of the wing upwash to
the calculated slope was only 0.02, so it is apparent that the esti-
mates of the upwash due to the fuselage must be charged with the major
portion of the discrepancy. This indicates that the potential-flow
equations as applied do not completely describe the flow around the
fuselage just outside the boundary layer at the location of the
detector.

To determine the effect of flap deflection, the increment of
indicated angle of attack produced by flap deflection is shown in
figure 4 as a function of flap deflection. Predicted results are
included for comparison. The agreement between the measured and the
predicted values is excellent. The effect of flap deflection is small
at the location of the detector (1.2 semispans forward of the quarter-
chord line) and could probably be neglected for most applications.
If the detector were closer to the wing,1 however, the change in upwash
produced by the flap deflection would become large enough to be sig-
nificant in fire-control use.


Flight Tests of F-86A-5


Figure 5 presents the comparison of the measured (ref. 1) and
predicted slopes and intercepts of the indicated versus true angle-of-
attack curves for the F-86A-5 airplane through the Mach number range.

In the case of a swept wing the spanwise position is also a powerful
factor in determining the amount of wing upwash, especially near the
wing (see ref. 2).

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NACA RM A53A15


Again the estimated slopes are higher than those actually measured.
It may be noted that the trends with increasing Mach number are in the
opposite direction. The theory also failed to predict the surprisingly
large variation of the intercept with Mach number. Since only 10 per-
cent of the calculated upwash is attributed to the wing (the detector
is 0.333 semispan forward of the quarter-chord line), it is apparent
that, again, the potential-flow equations used do not adequately
describe the flow at the location of the detectors which were outside
the boundary layer but near the surface of the fuselage.

In estimating the upwash angles, the fuselage was represented by
a semi-infinite body of revolution. It was assumed that there were no
effects of compressibility on the upwash due to the fuselage, and no
attempt was made to account for the air entering the inlet. The inlet-
flow and probable compressibility effects on the fuselage upwash may
account for a large portion of the discrepancy in slope shown in
figure 5. It is felt that the deviation of the actual fuselage from
the circular cross section used in the calculations combined with
compressibility effects is responsible for the poor correlation of the
measured intercept with that predicted.


CONCLUSIONS


Comparisons of predicted flow angles with those measured on the
forward portion of the fuselages of an aspect ratio 2 triangular-wing
wind-tunnel model and of an F-86A airplane in flight indicate the
following:

1. The accuracy of the predicted indicated angles of attack was
not sufficient to eliminate the necessity of a flight calibration of a
detector mounted on the fuselage. However, it did appear to be of
sufficient accuracy to be used for preliminary calculations in select-
ing the location of the detector and for the basic design of the
computer required to reduce the indicated to the true angle of attack.

2. The increase of upwash with flap deflection on the triangular-
wing wind-tunnel model was predicted accurately by the method described
herein.

3. The potential flow equations used to estimate the upwash due
to the presence of the fuselage did not completely describe the local
flow at the location of the detectors which were very close to the
fuselage but outside the boundary layer. The methods consistently
overestimated the indicated angle of attack.


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NACA RM A53A15


4. At the location of the detector on the fuselage of the wind-
tunnel model (1.2 semispans forward of the quarter-chord line), the
change in upwash due to flap deflection was small and could be
neglected for most applications.



Ames Aeronautical Laboratory
National Advisory Committee for Aeronautics
Moffett Field, Calif.


REFERENCES


1. McFadden, Norman M., Rathert, George A.,Jr., and Bray, Richard S.:
Flight Calibration of Angle-of-Attack and Sideslip Detectors on
the Fuselage of a 350 Swept-Wing Fighter Airplane.
NACA RM A52A04, 1952.

2. Rogallo, Vernon L.: Effects of Wing Sweep on the Upwash at the
Propeller Planes of Multiengine Airplanes. NACA TN 2795, 1952.

3. DeYoung, John, and Harper, Charles W.: Theoretical Symmetric Span
Loading at Subsonic Speeds for Wings Having Arbitrary Plan Form.
NACA Rep. 921, 1948.

4. Yaggy, Paul F.: A Method for Predicting the Upwash Angles Induced
at the Propeller Plane of a Combination of Bodies With an Unswept
Wing. NACA TN 2528, 1951.

5. Rogallo, Vernon L., McCloud, John L., III: Calculations of Upwash
in the Region Above or Below the Wing-Chord Planes of Swept-Back
Wing-Fuselage-Nacelle Combinations. NACA TN 2894, 1952.


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NACA RM A53A15


Note: All dimensions ore in feet.


12.50
25.00
23.91 --


Basic body of revolution used as body for

calculations. y = 2.25 f-(l-












Wing chord plane
8-_ ..- +






-.73 342r [4.322 K4.30! K4.00- L3.60 1 L 2.16

2.96 11.29 15.83 P706 36.43 42.30 50.2
Distance from nose of fusela/oge



Figure I.- Triongular-wing model used in 40-by 80-foot wind-
tunnel tests.


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NACA RM A53A15


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0 4 8
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Figure 2.- Variation of lift coefficient with geometric
ongle of attack for various flop deflections. Tri-
angular-wing model, 40 -by 80-foot wind tunnel.


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f 15.0


I






NACA RM A53A15


I
-I
-/--


0


/ 2 3 4 5 6 7 8 (, = o*)


Geometric angle of attack, a,, deg

Figure 3.- Indicated angle of attack as a function of geometric angle of
attack at various flap deflections. Triangular-wing model, 40-by 80-
foot wind tunnel.


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of change of indicated angle of attack with true angle
of attack and the indicated angle of attack at zero
true angle of attack for the F-86A-5 airplane.






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SECURITY INFQORIVIATON .6'i,.N
CONFIDENTIAL

UNIVERSITY OF FLORIDA

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