NASA Contractor Report 3311 

Rotary Balance Data for a 

Single-Engine Agricultural 

NASA 

CR 

3311 

c.1 

Airplane Configuration for an 

Angle-of-Attack Range of 80 to 900 

William Mulcay and Julio Chu 

CONTRACT NASl-14849 

DECEMBER 1980 
TECH LIBRARY KAFB, NM 

NASA Contractor Report 3311 

Rotary Balance Data for a 

Single-Engine Agricultural 

Airplane Configuration for an 

Angle-of-Attack Range of 80 to 900 

William Mulcay and Julio Chu 

Bihrfe Applied Research, Ix 

Jericho, New York 

Prepared for 

Langley Research Center 

under Contract NAS l- 14849 

NASA 

National Aeronautics 

and Space Administration 

Scientific and Technical 

Information Branch 

1980 
Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin 

tunnel are presented in plotted form for a l/lo-scale, singleengine agricultural airplane model. The configurations tested 

include the basic airplane, various wing leading-edge and wingtip devices, elevator, aileron and rudder control settings, and 

other modifications. Data are presented without analysis for an 

angle-of-attack range of 8O to 90°, and clockwise and counterclockwise rotations covering an g range from 0 to .9. This information is being obtained, along with that of other models, in 

order to establish a data base for analysis of the model and fullscale flight results, and to develop design charts for desirable 

stall/spin characteristics. 

INTRODUCTION 

A wind-tunnel rotary balance measures forces and moments 

acting on an airplane model while subjecting it to steady, though 

not necessarily equilibrium,spin flow conditions. The need for 

these data was recognized early in aircraft development. For example, early researchers noted in 1928 that the use of "straight" 

force tests conduce to erroneous conclusions when applied to the 

spin. NACA investigators at that time developed a "spinning" 

balance installed in a 5-foot vertical tunnel: they reasoned that 

subjecting all portions of an aircraft model to a constant wind 

velocity vector during static force tests was an unsatisfactory 

simulation of flow conditions existing in a spin. 

In recent years, NASA installed a rotary balance apparatus 

in the Langley full-scale tunnel and sponsored studies to determine the ability of current analytical techniques to compute the 
flight motions of current fighter configurations, Again, 

these studies clearly demonstrated the need to model the effects 

of spin rotation rate on the aerodynamic characteristics in all 

phases of the spinning motion. This concern is certainly warranted in the case of the high g (non-dimensionalized spin rate) 

values of today's general aviation airplane designs. 

These values are high due to high rates of rotation and low 

rates of descent for low wing-loaded vehicles. Hence, a difference of up to 90° in angle of attack between wing tips and 70° in 

sideslip angle between nose and tail can be realized. 

The NASA Langley Research Center has initiated a broad, general aviation stall/spin research program which includes spin 

tunnel and free-flight radio control model tests for a number of 

configurations typical of light, general aviation airplanes. To 

support these free-flight investigations, a rotary balance facility was developed in the Langley spin tunnel to rapidly identify 

an airplane's aerodynamic characteristics in a steady spin flow 

environment. The influence of spin radius, angle of attack, 

sideslip angle, rate of rotation and control settings on the 

aerodynamics can be measured, thereby permitting prediction of 

incipient and developed spin characteristics. 

A l/lo-scale, single-engine agricultural airplane model was 

tested on the wind tunnel rotary balance in the NASA Langley 

20-ft. spin tunnel. This report presents the data obtained for 

the basic configuration, various wing leading-edge and wing-tip 

devices, elevator, aileron and rudder control settings and other 

modifications. 

Use of trade names or names of manufacturers in this report does not 

constitute an official endorsement of such products or manufacturers, either 

expressed or implied, by the National Aeronautics and Space Administration. 

2 
SYMBOLS 

The units for physical quantities used herein are presented 

in the International System of Units (SI) and U.S. Customary 

Units. The measurements were made in the U.S. Customary Units: 

equivalent dimensions were determined by using the conversion 

factors given in reference 1. 

b 

E 

cL 

cN 

cA 

cY 

5 

'rn 

'n 

9 

S 

V 

a 

B 

R 

Rb 

Tm 

6a 

'e 

wing span, m (ft) 

mean aerodynamic chord, cm (in.) 

lift-force coefficient, Lift force 

qs 

normal-force coefficient, Normal force 

qs 

axial-force coefficient, Axial force 

qs 

side-force coefficient, Side force 

qs 

rolling moment coefficient, Rolling moment 

qSb 

Pitching moment pitching-moment coefficient, ~ qSE 

yawing-moment coefficient, Yawing moment 

qSb 

free-stream dynamic pressure, X/m2 (lb/ft') 

wing area, m2 (ft2 1 

free-stream velocity, m/set (ft/sec) 

angle of attack, deg 

angle of sideslip, deg 

angular velocity about spin axis, rad/sec 

spin coefficient, positive for clockwise spin 

aileron deflection, positive when right aileron is down 

(6 -6 

aright aleft 

l/2, deco 

elevator deflection, positive when trailing edge is down, 

deg 

rudder deflection, positive when trailing edge is to left, 

de-g 
Abbreviations: 

cg center of gravity 

LE leading edge 

SR spin radius 

TEST EQUIPMENT 

A photograph and sketch of the rotary balance apparatus installed in the Langley 20-Foot Spin Tunnel are shown in figures 

1 and 2, respectively. The arm of the balance system rotates 

about a vertical axis and is mounted on the free end of a horizontal supporting boom, which is hinged at the wall. It is moved 

from the wall to the center of the tunnel by cables, secured, and 

then driven by a drive shaft through couplings and gears. 

A test model is mounted on a strain gauge balance which 

is affixed to the bottom of the rotary balance apparatus. Controls located outside the tunnel are used to activate motors on 

the rig, which position the model to the desired attitude. The 

angle-of-attack range of the rig is 8 to 90 degrees, and the 

sideslip angle range is 215 degrees. The spin radius and lateral 

displacement motors allow the operator to position the moment 

center of the balance on the spin axis or at a specific distance 

from the spin axis. This is done for each combination of angle 

of attack and sideslip angle. The NASA six-component strain 

gauge balance mounted inside the model measures the normal, axial, 

and side forces, and the yawing, rolling, and pitching moments 

acting about the model body axis. Interactions among the six 

components are available from previous laboratory calibration 

tests and are used to correct the raw forces and moments output 

by the balance strain gauge transducers. For the data presented 

herein and the practice generally, the force and moment axes of 

the balance are mounted coincident with the body axes of the air- 
plane, with the moment center at the cg location about which 

the aerodynamic moments are desired, Electrical current to 

and from the balance and to the positioning motors on the rig, 

and position-indicating potentiometer voltages are conducted 

through slip rings located at the rig head. Examples of how 

the rig is positioned for different angles of attack and sideslip angles are shown in figures 2a and 2b, respectively. 

The model can be rotated up to 90 rpm in either direction. 

By using different rotational speeds and a specific airflow in 

the tunnel, the motions of an airplane in a steady spin can be 

simulated. The aerodynamic forces and moments can be measured 

Rb at different values of 2v, including the case of g = 0, where 

static aerodynamic data are obtained. 

Additional background for this apparatus is given in 

reference 2. 

The data acquisition, reduction‘and presentation system for 

the rotary balance set-up includes an RPM indicator, a 12-channel 

scanner/voltmeter with a 14-bit variable gain analog to digital 

converter, and a mini-computer with graphics and a printer. With 

this equipment, immediate presentation of data in tabular and/or 

plotted form is possible. 

TEST PROCEDURES 

Rotary aerodynamic data are obtained in two steps. First, 

the inertial forces and moments (tares) acting on the model at 

different attitudes and rotational speeds are measured. To accomplish this, a covered bird cage like structure, which encloses 

the model without touching it, is mounted on the upper portion 

of the rig and rotates with the model. In this manner, the air 

immediately surrounding the model is trapped by the structure 

and constrained to rotate with it. As the rig is rotated at the 

desired attitude and rate, a pulse voltage from a magnetic 

5 
trigger device on the drive shaft begins the data sampling 

process. Eighty samples of the 6 balance output channels are 

taken over 10 revolutions, with the computer adjusting the 

timing, according to the RPM indicator, to sample at 8 fixed 

angular positions of the rotating rig. The inertial forces 

and moments thus measured are curve-fitted to a least-squares 

R2 function of loo and stored for later use on magnetic tape. 

For the second step, the covered structure enclosing the 

model is removed and the total wind-on forces and moments are 

measured. The aerodynamic forces and moments are determined by 

subtracting the tare values from the total wind-on values at 

each tested RPM and angle of attack. The aerodynamic data are 

converted to coefficient form and stored on magnetic tape. 

MODEL 

A l/lo-scale fiberglass-epoxy, wood and aluminum model of 

a typical single-engine configuration used for agricultural purposes was tested in the present study. A three-view drawing of 

this model is shown in figure 3, dimensional characteristics of 

the model are presented in Table I, and a photograph of the model 

installed on the rotary balance located in the Langley 20-Foot 

Spin Tunnel is presented in figure 1. 

The model was fabricated such that various modifications 

could be added. These consisted of a turbine nose, as shown in 

figure 4; canopy fairing and a wing-fuselage fillet, shown in 

figure 5; an outboard leading-edge wing droop, shown in figure 6; 

and a full span leading-edge wing slot, shown in figure 7. In 

addition, several wing-tip modifications designed to alter the 

characteristics of wing-tip vortex flow were tested. These 

consisted of winglets, wing-tip sails and Lockheed Vortex Diffuser 

Vanes, as shown in figures 8 and 9. The Lockheed Vortex Diffuser 

Vane is patented by the Lockheed-Georgia Aircraft Company. 
The model control surfaces could be fixed at any position 

prior to the test, The maximum deflections for the control surfaces were: 

Elevator, deg 27 UP, 17 down 

Rudder, deg 24 right, 24 left 

Aileron, deg 21 up, 17 down 

TEST CONDITIONS 

The tests were conducted in the spin tunnel at a tunnel velocity of 7.62 m/set (25ft/sec), which corresponds to a Reynolds 

number of approximately 120,000 based on the model mean aerodynamic chord. Unless otherwise noted in Table II, all the configurations were tested through an angle-of-attack range of 8 to 

35O, with the spin axis in the plane of symmetry of the aircraft 

at a full-scale distance 152.4 cm (60 in.) forward of the cg 

location of .26e. In addition, for some configurations, data 

were obtained for angles of attack between 30 and 90°, with the 

spin axis through the cg location- Data were thus obtained at 

two spin radii for angles of attack of 30 and 35 degrees. At 

each spin attitude, measurements were obtained for nominal !$ 

values of .l, -2, .3, -4, .5, -6, -7, .8, and .9, in both clockwise and counter-clockwise directions, as well as for g = 0 

(static value). 

DATA PRESENTATION 

Table II identifies the configurations tested and the 

corresponding appendix figure numbers which present the aerodynamic data. The aerodynamic coefficients plotted vs. 22 2vare 

presented for each configuration in six sequentially numbered 

figures in the following order: Cn, CR, Cm, CN, Cyr and CA. 

Each figure consists of two pages, which present the subject 
aerodynamic coefficient plotted vs g for the following angles 

of attack and spin radius: 

al a = 8, 10, 12, 14, 16 deg SR = 152.4 cm (60 in.) 

b) cj = 18, 20, 25, 30, 35 deg SR = 152.4 cm (60 in.) 

and, in addition, for some configurations noted in Table II: 

cl cx = 30, 35, 40, 45, 50 deg SR = 0 

d) a = 55, 60, 70, 80, 90 deg SR = 0 

All the moment data are presented for a cg location of .26C. 

Lift coefficient plotted vs angle of attack for zero rotation rate is presented in the Appendix for several configurations cited in Table II. 

8 
REFERENCES 

1. Mechtly, E.A.: The International System of Units - Physical 

Constants and Conversion Factors. NASA SP-7012, 1973; 

2. Bihrle, William, Jr.; Hultberg, Randy S.; and Mulcay, 

William: Rotary Balance Data for a Typical Single-Engine 

Low-Wing General Aviation Design for an Angle-of-Attack 

Range of 30° to 90°. NASA CR-2972, 1978. 
TABLE I.- DIMENSIONAL CHARACTERISTICS OF TI;E BASIC MODEL 

Overall length, m (ft) . . . . . . . . . . . . . .89 (2.92) 

Wing: 

Span, m (it) . . . . . . . . . . . . . 

Area, m2 (ft2) . . . . . . . . . . . . 

Root chord, cm (in.). . . . . . . . . . 

Tip chord, cm (in.) . . . . . . . . . . 

Mean aerodynamic chord, cm (in.). . . . 

Aspect ratio . . . . . . . . . . . . 

Dihedral, deg . . . . . . . . . . . . 

Incidence: 

Root, deg 

Tip (of constant chord section), deg 

Airfoilsection . . . . . . . . . . . 

Horizontal tail: 

Span, m (it) . . . . . . . . . . . . 

Area,m2 (it21 . . . . . . . . . . . . 

Incidence, deg . . . . . . . . . . . . 

Airfoil Section . . . - . - - - - - - - 

Vertical tail: 

. . . . 49 (1.60) 

. . . . 050 (.54) 

. . . -1.55 

. . . Flat plate 

Area, m2 (ft2) . . . . . - . - . - . . -021 (.222) 

Airfoil section Flat plate 

. . . 

. . . 

. . . 

. . . 

. . . 

. . . 

. . . 

. . . 

1.35 (4.43) 

. 304 (3.27) 

22.9 (9.0) 

22.9 (9.0) 

22.9 (9.0) 

6.0 

3.5 

2.07 

0.39 

NACA 4412 

10 
TABLE II.- CONFIGURATIONS TESTED AND FIGURE INDEX 

(Unless noted otherwise, all configurations tested through a=8 to 35' at SR=152.4 cm (60 in.) and fi=O'.) 

FIGURE I CONFIGURATION I I 

I REMARKS 

NO 

; Ae j Aa j br 

I / deg 1 deg j deg 

I I 

a,b,c Al-A6 

I 

/Basic 

: 

0 

I 17 

! -27 i ! 

0 also a=30 to 90' at SR=O 

A7-A12 -24 

A12-A18 

A19-A24 

A25-A30 

A31-A36 

A37-A42 

'A43-A48 

'A49-A54 

'A55-A60 

aA61-A66 

aA67-A72 

b A73-A78 

b A79-A84 

b A85-A90 

i v 

! 

i 1 

jTurbine nose i 0 

/Canopy fairing 

! 1 

iWing-fuselage fillet 

; ; / 

1 

4 

with outboard LE wing droop ! I 

i v 

with outboard LE wing slat 

;Full-span wing slat I 

/Segmented wing slat 

Winglets 

Wing tip sails 

'Lockheed Vortex Diffuser Vanes I 

I 

+19 

-19 

0 

I 

A also a=30 to 90' at SR=O 

also a=30 to 90' at SR=O 

a CL ys. CY. presented in figure A91. 

b CL vs. c1 presented in figure A92. 

P P c CL vs. c1 presented in figure A93. 
Figure 1. - l/l0 - scale model instal led on rotary balance apparatus. 
A Slip ring housing 

B Spin radius offset 

potentiometer 

C Lateral offset 

drive gears 

D Lateral offset 

potentiometer 

E Strut 

F Sideslip angle 

potentiometer 

G- Sideslio anole 

positioning-motor 

e 

-C 

i4 

i 

E 

Spin axis 

Velocity vector 

I 

(b) Front view of model. 

Figure 2.- Concluded. I3 

I I 
A Slip ring housing 

B Drive shaft 

C Support boom 

D Spin radius offset 

potentiometer 

E Counterweight 

F Strut 

G Angle of attack 

positioning motor 

Spin axis 

Velocity vector t 

(a) Side view of model. 

Figure 2.- Sketch of rotary balance apparatus. 

14 
/!A 4.72 V. Jd 

(1. i36’ $2.50) 

I 

J 

t 

22.9 

(9.,0) I 

e 

1 

u --I 

-u 

uselage reference line __ 2 

Figure 3. -Three-view of l/l0 - scale basic configuration. Center of gravity positided at 

0.26L Dimensions are given in centimeters (inches), model scale. 

is 
.56 

u 0 

c 22) 

Figure 4. -Turbine nose tested on model, 

model scale. 

Dimensions are given in centimeters (inches), 
Figure 5. - Canopy fairing and wing-fuselage fillet tested on model. Dimensions are given 

in centimeters(inches1, model scale. 

I7 
m 

I 

t 

- Fuselage centerline 

Wing chotd line./ 

/ 

J: 

r---- ------------1 

---- __--- -- --- 

P,:,-$- "u-i 

\ 

Figure 6. - Wing-fuselage fillet with outboard LE wing droop tested on model. Dimensions 

are given in centimeters (inches), model scale. 

I 
tFuselage centerline 

Wing chord line - 

- 13.3 

(5.25) -I--- 

13.3 - 

(5.25) - (K) ----I 

CyQur NACA 4412 

“: 

Contour GA(W)-1 I’

F&ire 7. - Winpfuselage fillet and LE wing slat configurations tested on model. Dimensions are given 

in centimeters(inches1, model scale. 
---- 

------- 

I 

Modified GA(W)-? cm+inn 

Figure 8. - Winglet installation tested on model. -Dimensions are given in centimeters 

model scale. 

20 
9.37 

(3.69) 

L68 

f.66) 

2.74 

(1.08) 

Right wing tip sail 

Lockheed vortex diffuser vane 

Figure 9. - Wingtip sails and tip vortex diffusers tested on model. Dimensions are given 

in centimeters (inches), model scale. 
APPENDIX 

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